3 1176 00156 1985 - ntrs.nasa.gov
Transcript of 3 1176 00156 1985 - ntrs.nasa.gov
3 1176 00156 1985" NASA-CR-144887
1.9800006814NASA Contractor Report 144887
•3
,: ANALYSES AND TESTS OF THE B-1 AIRCRAFT
STRUCTURAL MODE CONTROL SYSTEM
John H. Wykes, Thomas R. Byar, Cary J. MacMiller,and David C. Greek
Contract NAS4-2519 '-January 1980
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CR-144,887
Wgkes, d. H.Analgses and tests o_ the B-1 aircraft structuralmode control system.NASA
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NASA Contractor Report 144887
ANALYSES AND TESTS OF THE B-1 AIRCRAFT
STRUCTURAL MODE CONTROL SYSTEM
John H. Wykes, Thomas R. Byar, Cary J. MaeMiller,and David C. GreekRockwell InternationalEl Segundo, California
Prepared forDryden Flight Research Centerunder Contract NAS4-2519
NIA ivvo- i 73NationalAeronauticsandSpaceAdministration
1980
TABLE OF CONTENTS
Page
SU_tIARY I
IKI'RODUCTION 1
FLEXIBLEAIRCRAFT EQUATIONSOF MOTION 3
FLEXIBLEAIRCRAFTANALYSESMODEL 22
DynamicAnalysis System 22Free-FreeVibrationModal Data 24
AerodynamicData 27Control-SurfacesInertiaReactionForces 38
Active ControlSystems 42
COMPARISONSOF ANALYSESAND FLIGHT-TESTRESULTS 42
FORWARDSMCS SENSOR PACKAGERELOCATION 58
TRUNCATEDANALYTICALMODELS 78
ANALYSISOF SNCS VANEAERODYI_.IICINTERFERENCEEFFECT 90
IMPACTOF SHCSON SELECTEDLOADS 101
Backgroundof Using SMCS in the B-I FatigueAnalysis I01Gust Loads Analysis Description 102Gust Loads StructuralModel 103"
Gust Loads Aerodynamics 104Gust Loads ControlSystems 106Load Method Discussion 106Gust StatisticalLoad Calculations II0
Load Phasing IIi
ConditionMatching 112-_ ExampleConditionLoad Results 113
SMCS VANE EFFECT ON INLET/ENGINECHARACTERISTICS 154
Test Description 155
Flight-TestInstrumentation 159
i
TABLE OF CONTENTS - (Continued)
Page
Test Results 165
OscillatingVanes 167StaticVane Deflections 170
Maneuvers 170
Engine ThrottleTransients 195
Summary 205
SUMMARYOF SMCS FLIGHT TEST RESULTS 205
SMCS Performancein Time-History-DataFormat 206SMCS Performancein PSD-Data Format 213
SMCS and HandlingQualities 213SMCS High-GainTests 218Crew Evaluationsof SMCS Effects 220
OverallRide Quality 220Flight-PathControlTasks 225Non-Flight-PathControlTasks 225Readabilityof Instrumentsand Displays 225Reaching/UsingControls 226Crew Fatigue 227Motion Sickness 227
PhysicalDiscomfort 228AdditionalRide QualityObservations 228Handling Qualities 229
APPENDIX - NOMENCLATURE. 230
General 230
Ride QualityEquationsof Motion Related 234Load Equationsof Motion Related 247Engine/InletRelated 250
REFERENCES 253
ii
LIST OF ILLUSTRATIONS
Figure Title Page
1 B-I aircraftwith wings swept aft. 22 Sign conventionfor rotationaland linear rates and
accelerations. 15
3 Sign conventionfor coefficientsand control surfacedeflections. 16
4 Structuralmode deflections,slopes,and generalizedforces sign convention. 17
5 Angle-of-attack(_) and sideslip QS) definitions. 196 Euler angle definitionsand rotationsequence. 207 Dynamicanalysissystem. 238 Typical elasticaxes and mass point locations. 259 Fuselageelasticaxis refinement. 26i0 Typical symmetricmode vector plot. 29ii Typical antisymmetricmode vector plot. 3012 Typical structuralmode deflectiongrid points. 3113 Panellingand box grid for Doublet Latticeaerodynamics
program. 3214 Typical low-frequencyadjustmentmade to analytical
frequency-dependentaerodynamicdata. 3515 Typical aeroelasticflexible-to-rigidratio data for
aerodynamiccoefficientsas a functionofparticipatingstructuralmodes. 39
16 Typicalcontrolsurface inertiareactiongeneralizedforces. 41
17 Pitch axis SCAS analyticalmodel. 4318 Yaw axis SCAS analyticalmodel. 4419 Roll axis SCAS analyticalmodel. 45
20 Vertical SMCS analyticalmodel. 4621 LateralSMCS analyticalmodel. 4722 Comparisonof flight test and analyticaldata,
frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off. 49
23 Comparisonof flight test and analyticaldata,frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS vane deflection,
SCAS on, SMCS off. 5024 Comparisonof flight test and analyticaldata,
frequencyresponseof normal load factor atFS 571.5 (225)due to SMCS deflection,SCASon, SMCS on. 51
iii
LIST OF ILLUSTRATIONS- (Continued)
Figure Title Page
25 Comparisonof analyticalSMCS actuatormodels toflightand simulatortest data. 53
26 Effect of test-derivedSMCS actuatormodel, frequencyresponseof normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on. 54
"'27 Comparisonof flight test and analyticaldata,frequencyresponseof lateralload factor at FS 571.5(225)due to SMCS vane deflection,SCAS off, SMCS off. 55
28 Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 56
29 Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. 57
30 Effect of SMCS lateralgain on power spectraldensityof lateral load factor at pilot stationFS 746.8 (294). 59
31 SMCS sensorpackage locationsand couplingcharacteristics. 60
32 Effect of SMCS sensor package locationon lateralloadfactor at FS 571.5 (225)due to differentialvanedeflection,analyticaldata. 61
33 First fuselageverticalbendingmode dampingversusSMCS gains for originaland relocatedforwardsensorpackage. 63
34 SMCS vertical axis performancewith relocatedforwardsensorpackage, frequencyresponseof normal loadfactor at FS 515.6 (203)due to SMCS vane deflection,case I. 65
35 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 515.6 (203)due to SMCS differentialvane deflection,case i. 66
36 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 746.8 (294)due to SMCS differentialvane deflection,case I. 67
37 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateral loadfactor at FS 515.6 (203)due to SMCS differentialvane deflection,case 2. 68
iv
LIST OF ILLUSTRATIONS - (Continued)
Figure Title Page
38 SMCS lateralaxis performancewith relocatedforwardsensor package,frequencyresponseof lateralloadfactorat FS 746.8 (294)due to SMCS differentialvane deflection,case 2. 69
39 SMCS verticalaxis performancewith relocatedforward
sensorpackage, frequencyresponseof normal loadfactor at FS 515.6 (203)due to SMCS vanedeflection,case 2. 70
40 SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralloadfactor at FS 515.6 (203)due to SMCS differential
vane deflection,case 3. 7241 SMCS performancewith relocatedsensor,PSD of vertical
load factor at FS 746.8 (294) - pilot station,case i. 7342 SMCS performancewith relocatedsensor,PSD of lateral
load factor at FS 746.8 (294) - Pilot station,case i. 74
43 SMCS performancewith relocatedsensor PSD of verticalload factorat FS 746.8 (294) - pilot station,case 2. 75
44 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6 (203) sensor location,case I. 76
45 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) pilot station,case 2. 77
46 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6 (203) - sensor location,case 2. 79
47 SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) - pilot station,case 3. 80
48 SMCS performancewith relocatedsensor,PSD of lateralload factorat FS 515.6 (203) - sensor location,case 3. 81
49 SMCS performancewith relocatedsensor,PSD of lateralload factorat FS 746.8 (294) - pilot station,case 4. 82
50 Normal load factordue to horizontaltail frequency
response comparisonsof full and truncateddynamicanalyticalmodel, SCAS on. 91
51 Typicalvane-inducedinterferenceforces. 9352 Effect of SMCS vane aerodynamicinterferences,
frequencyresponse of normal load factor at FS 571.5(225) due to SMCS vane deflection,SCAS off, SMCS off. 95
V .
LIST OF ILLUSTRATIONS (Continued)
Figure Title Page
53 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 96
54 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. 97
55 Effect of SMCS vane aerodynamicinterferences,frequency
responseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS off, SMCS off. 98
_6 Effect of SMCS vane aerodynamicinterferences,frequency
responseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS off. 99
57 Effect of SMCS vane aerodynamicinterferences,frequencyresponseof lateralload factorat FS 571.5 (225)dueto SMCS vane deflection,SCAS on, SMCS on. I00
58 Loads analysis SIC point locationsand DoubletLatticegeometry. 105
59 Controlsystemsfor gust loads. 10760 Structuralstationsfor gust loads analysis. 10861 Load calculationequations. 10962 SMCS off, wing bendingmoment frequencyresponse -
WS 985 cm (387.6in). 11563 SMCS off, wing bendingmoment responsepower spectrum
WS 985 cm (387.6in.). 11664 SMCS off, wing bendingmoment exceedances-
WS 985 cm (387.6in.). 11765 SMCS off, forebodybendingmoment frequencyresponse
FS 1377 cm (542in.). 11866 SMCS off, forebodybendingmomentresponse power
spectrum - FS 1377 cm (542 in.). 11967 SMCS off, forebodybendingmoment exceedances
FS 1377 cm (542 in.). 12068 SMCS off, forebodybendingmoment frequencyresponse
FS 2367 cm (932 in.). 12169 SMCS off, forebodybendingmoment responsepower
spectrum - FS 2367 cm (932 in.). 12270 SMCS off, forebodybendingmoment exceedances-
FS 2367 cm (932 in.). 12371 SMCS off, normal load factor frequencyresponse -
aircraftCG. 124
vi
LIST OF ILLUSTRATIONS(Continued)
Figure Title Page
72 SMCS off, normal load factor responsepower spectrum -aircraftCG. 125
73 SMCS off, normal load factor exceedances aircraftCG. 12674 SMCS off, normal load factor frequencyresponse -
pilot station. 12775 SMCS off, normal load factor responsepower spectrum-
pilot station. 12876 SMCS off, normal load factor exceedances pilot station. 12977 SMCS off, delta horizontalstabilizerfrequencyresponse. 13078 SMCS off, delta horizontalstabilizerresponsepower
spectrum. 13179 SMCS off, delta horizontalstabilizerexceedances. 13280 SMCS on, wing bendingmoment frequencyresponse -
WS 985 cm (387.6in.). 133
81 SMCS on, wing bendingmoment responsepower spectrum -WS 985 cm (387.6in.). 134
82 SMCS on, wing bendingmoment exceedances- FS 985 cm(387.6in.). 135
83 SMCS on, forebodybendingmoment frequencyresponse-FS 1377 cm (542 in.). 136
84 SMCS on, forebodybending moment responsepower spectrum -FS 1377 cm (542 in.). 137
85 SMCS on, forebodybendingmoment exceedances FS 1377 cm(542 in.). 138
86 SMCS on,forebody bendingmoment frequencyresponseFS 2367 cm (932in.). 139
87 SMCS on, forebodybendingmoment responsepower spectrum -FS 2367 cm (932in.). 140
88 SMCS on, forebodybendingmoment exceedances-FS 2367 cm (932in.). 141
89 SMCS on, normal load factor frequencyresponse -aircraftCG. 142
90 SMCS on, normal load factor responsepower spectrum-aircraftCG. 143
91 SMCS on, normal load factorexceedances- aircraftCG. 14492 SMCS on, normal load factor frequencyresponse -
pilot station. 14593 SMCS on, normal load factorresponsepower spectrum -
pilot station. 14694 SMCS on, normal load factor exceedances- pilot station. 14795 SMCS on, delta horizontalstabilizerfrequencyresponse. 14896 SMCS on, delta horizontalstabilizerresponsepower
spectrum. 149
vii
I,IST OF ILLUSTRATIONS - Continued
Figure Title Page
97 SMCS on, delta horizontalstabilizerexceedances. 15098 SMCS on, delta mode controlvane frequencyresponse. 15199 SMCS on, delta mode controlvane responsepower spectrum. 152i00 SMCS on, delta mode controlvane exceedances. 153I01 B-I flight test aircraft,wings swept 65 degrees. 156I02 B-I air inductionsystem. 157103 B-I inlet subsonicdiffuser flow area distribution. 158
104 B-I inlet/engineaerodynamicinterfaceplaneinstrumentation. 160
105 B-I flight-testAIP data acquisitionand signalconditioningsystem. 161
106 In-flightcalibrationcycle, total pressuresataerodynamicinterfaceplane, flight 1-5, 64 sps. 163
I07 Representativediscriminatoroutput,AIP total-pressureCBW data, flight i-i0. 164
108 AIS/SMCS investigation- flight 2-19,Mach 0.83. 168I09 AIS/SMCS investigationflight 2-19,Mach 0.83/16,000feet,
= 3 degrees,vortex ingestionin no. 2 inlet during
sideslipoperation,PLA = max. 169ll0 AIS/SMCS investigation,flight 2-19, _ = 3 degrees,
vortex ingestionin no. 1 inlet during sideslipoperation. 171
Iii AIS/SMCS investigation,flight 2-19, ff= 3 degrees,vortex ingestionin no. 2 inlet during IDLE-INTthrottle transient. 172
If2 Flight 2-33 test matrix and AIP signalsduringMach 0.85operationwith SMCS vanes deflected20 degrees,no. 2 inlet,RB = 7 degrees,RC = 5 degrees 173
113 No. 2 inlet,effect of sideslipangle on steady-stateinlet characteristicsat SMCV= 20 degrees, flight 2-33. 174
114 No. 2 inlet,effect of sideslipangle on steady-stateinlet characteristicsat various SMC vane angles,flight 2-33. 175 .
115 Time historiesof total-pressurerecoveryand stall-margin index,Mach 0.85, with SMCS vane deflectedand 0 and 20 degrees,no. 2 inlet, RB = 7 degrees,RC = 5 degrees,flight 2-33. 177
116 Dynamiccircumferentialand radial distortioncomponents,Mach 0.85, with SMCS vane deflected0 and 20 degrees,no. 2 inlet, RB = 7 degrees,flight 2-33. 178
viii
LIST OF ILLUSTRATIONS - (Continued)
Figure Title Page
117 Dynamic total pressurecontoursduringMach 0.85 179operationWith SMCS Vane deflected0 and 20degrees,no. 2 inlet, RB = 7 degrees,RC = 5de_rees, flight 2-33.........
118 Effectsof SMC vane positionon no. 2 inlet steady-stateand dynamic characteristics,Mach = 0.85 and alpha =
" 2.7 degrees,flight 2-37. 180119 Effects of sideslipangle on steady-stateinlet
characteristicswith SMCV= 0 degreesand alpha = 1
degree,flight 2-37. 181120 Effectsof sideslipangle on steady-stateinlet
characteristicswith SMCV = 20 degreesand alpha =
1 degree, flight= 2-37. 182121 Effectsof SMCVwake on no. 2 inlet dynamictotal-
pressurecontoursduring sideslip,SMCV = 20 degreesand alpha = 1.4 degrees,flight 2-37. 183
122 Effectsof sideslip angle on steady-stateinletcharacteristicswith SMCV = -8 degrees and alpha =1 degree,flight 2-37. 184
123 Effectsof SMCV wake on no. 2 inlet dynamictotal-
pressure contoursduring sideslip,SMCV = -8 degreesand alpha = 0.9 degrees,flight 2-37. 185
124 Effectsof sideslipangle on Steady-stateinletcharacteristicswith SMCV= 0 degreesand alpha =2.6 degrees,flight 2-37. 186
125 Effectsof sideslipangle on steady-stateinletcharacteristicswith S_4CV= 20 degreesand
alpha = 2.6 degrees,flight 2-37. 187126 Effectsof SMCVwake on no. 2 inlet dynamictotal-
pressure contoursduring sideslip,SMCV = 20 degreesand alpha = 2.6 degrees, flight 2-37. _ 188
127 Effects of sideslipangle on steady-stateinlet1 characteristicswith SMCV = 0 degreesand
alpha = 5.8 degrees,flight 2-37. 189_ 128 Effectsof sideslipangie on steady-stateinlet
characteristicswith SMCV = 20 degreesand
alpha = 5.8 degrees,flight 2-37. 190129 IndividualAIP probes, total-pressurerecoveryarray,
SMCS vane deflected20 degrees,Mach 0.85, _ = 6 degrees,flight 2-37. 191
130 Variations in circumferentialand radial distortion
componentsduring sideslipoperationwith the SMCS vanedeflected+20 degrees,Mach 0°85, _ = 6 degrees,
flight 2-37. 192
ix
I,IST OF ILLUSTRATIONS- (Continued)
F i gure Title Page
131 Effectsof SMCV wake on no. 2 inlet dynamictotal-pressure contoursduring sideslip,SMCV= 20 degreesand alpha = 5.8 degrees,flight 2-37. 193
132 Effectsof SMCVwake on no. 1 inlet dynamictotal-pressure contoursduring sideslip,SMCV = 20degreesand alpha = 5.8 degrees,flight 2-37. 194
133 Effectsof angle-of-attackmaneuverson steady-stateinlet characteristicswith SMC vane at 20 degrees,flight 2-42. 196
134 Effectsof an angle-of-attackmaneuver on steady-stateinlet characteristicswith no. 1 engine at intermediate,flight 2-42, SMCS vane deflectionangle = 0 degrees. 197
135 Effectsof an angle-of-attackmaneuveron steady-stateinlet characteristicswith no. 2 engine at intermediate,flight 2-42, SMCS vane deflectionangle = 0 degrees. 198
136 Effectsof an angle-of-attackmaneuver on steady-stateinlet characteristicswith no. 2 engine at intermediate,flight 2-42, SMCS vane deflection angle = 0 degrees, 199
137 Effectsof sideslipangle on steady-stateinletcharacteristicswith SMCV= 13 degreesandalpha = 3.3 degrees,flight 2-38. 200
138 Effectsof SMCVwake on no. 2 inlet dynamictotal-pressure contoursduring sideslip,SMCV = 13 degreesand alpha = 3.3 degrees, flight 2-38. 201
139 Steady-'stateinlet characteristicsduring no. 2 enginetransients,SMCV = 20 degrees, alpha = 5.5 degrees,and beta = +4 degrees, flight 2-38. 202
140 Steady-stateinlet characteristicsduring no. 2 enginetransients,SMCV= 20 degrees, alpha = 3 degrees, andbeta = +4 degrees,flight 2-38. 203
141 Steady-stateinlet characteristicsduring no. 2 enginetransients,SMCV= 20 degreesand -8 degrees,alpha = 1degree with positive sideslip,flight 2-38. 204
142 Typical dynamicresponsenear crew stationdue to
turbulenceduring low-altitude,high-speedflight. 207143 SMCS performancein turbulenceM = 0.70 alt = 305 m
(I,000ft) AGLA= 65°. 208
144 First fuselagesymmetricstructuralmode dampingfrom horizontaltail pulse excitations. 210
X
LIST OF ILLUSTRATIONS- (Concluded)
Figure Title Page
145 Effect of SMCS verticalgain settingon first fuselagesymmetricstructuralmode dampingat variousaircraftweights. 211
.- 146 First fuselageantisymmetricstructuralmode dampingfrom forced SMCS vane oscillations. 212
147 Vertical SMCS performancein turbulenceas shown bypower spectraldensity of vertical load factor atpilot station,FS 747 (294). 214
148 Lateral SMCS performancein turbulenceas shown bypower spectraldensityof lateralload factor atpilot station,FS 747 (294),single-peakresponse. 215
149 Lateral SMCS performancein turbulenceas shown bypower spectraldensityof lateralload factor atpilot station,FS 747 (294),double-peakresponse. 216
150 SMCS impact on short-periodand dutch roll frequencies. 217151 Ride qualityratingsfor varyingdegreesof turbulence. 221152 Effectsof turbulenceon flightpath controltasks. 221153 Effectsof turbulenceon tasks other than flightpath
control. 222
154 Effectsof turbulenceon readabilityof instrumentsand displays. 222
155 Effectsof turbulenceon reachlng/usingcontrols. 223156 Effectsof turbulenceon crew fatigue. 223157 Effectsof turbulenceon tendencyfor motion sickness. 224158 Effectsof turbulenceon physical discomfort. 224
xi
LIST OF TABLES
Table No. Title Page
I GeneralFlexibleVehicle Equationsof Motion, Time Domain 5II Equationsof Motion for LongitudinalRigid-Bodyand
SymmetricStructuralModes, FrequencyDomain I0I[l Equationsof Motion for Lateral-Directional,Rigid-Body
and AntisymmetricStructuralModes, FrequencyDomain 12IV TypicalAnalyticalStructuralMode Data 28V Longitudinal-SymmetricAero CoefficientsFrom Doublet
LatticeProgram 37VI SMCS Vane EffectivenessIncludingInterferenceEffects,
Longitudinal-SymmetricCase, FrequencyDomain 94VII Gust Loads Model StructuralDegreesof Freedom 103VIII Load Comparison,SMCS Off Versus SMCS On 114IX SMCS Vane Effectson Inlet/EngineCharacteristics,
FlightTest Investigation 166X SMCS High-GainTests Summary 219
xii
ANALYSESAND TESTS OF THE B-I AIRCRAFT STRUCrURALMODE CONTROLSYSTIg4
John H. Wykes, Thomas R. Byar, Cary J. MacMiller,and David C. Greek
Rockwell International,North AmericanAircraftDivision
- _y
An 18-monthprogramwas conductedto compileand document for publicationinformationpertainingto analysesand flight tests of the B-I StructuralModeControlSystem (SMCS). This is the secondphase of a continuingeffort;results from the first phase study are documentedin referencei. This reportcovers the followingtopics:
(I) Flexible aircraftequationsof motion
(2) Descriptionof flexibleaircraftanalysesmodel
(3) Comparisonof analysesand flight-testperformanceresults of theSMCS
(4) A stmmaryof the study of the forward_4CS sensorpackagerelocation
(5) Truncatedanalyticalmodels used in simulationeffort
(6) An analysisof the SMCSvane interferenceeffects
(7) ImFactof SMCS on selectedloads
(8) Flight-testresultsof the SMCS vane effectson inlet/enginecharacteristics
" - (9) Stmmaryof SMCS flight-testresults
INTROII/CTION
The B-I aircraftis one of the first vehiclesto includea controlcon-
figuredvehicle (CCV) conceptin the early design phases. The aircrafthas arequirementto providea specifiedlevel of ride quality for the crew. Thisrequirementhas been met on the B-I throughthe use of an automaticcontrolsystem (_CS) whose main externalfeatureis a set of vanes (near the crewstation)which are canted down 30 degreesfrom the horizontal. (See figurei.)
N
Structural modecont ro I vanes
Figure 1. - B-1 aircraft with wings swept aft.
A sM_stantialsavingsinweightwas achievedwith thisapproachas comparedto direct material stiffening. The details of systom re_luiranents had to bedetermined from a production (long-life) point of view, which has not beendone before for a system of this type. Extensive wind tunnel tests of thevane characteristics were conducted. Analytical models of the flexible air-craft and control systems were developed to analyze requirements and toinvestigate stability and performance. Component parts were tested to the
.- requirements in the laboratory. Flight tests of the SMCShave been conducted,and comparisons with analytical predictions have been made. Because of all ofthis, it has been recognized that the B-1 offers an excellent opportunity formuch needed further evaluation of such a system as the SMCSto insure theoptimun use of these systems for future applications.
The overall objective of this research area is to compile and documentinformation about the conceptual design, development, and flight tests of theB-1 SMCSand its impact on ride quality. Since the B-1 is the first aircraftto have a system such as the SMCSdesigned for production and long-service use,it is expected that the reports prepared will add to the technology base fordesign of future large military or civil aircraft. The specific overallobjectives are to:
(1) Investigate the improvements in total dynamic response of a flexibleaircraft and the potential benefits to ride qualities, handling qualities,crew efficiency, and reduced dynamic loads on the primary structures
(2) Evaluate the effectiveness and performance o£ the SMCS,which usessnell aerodynamic surfaces at the vehicle nose to provide damping to thestructural modes
The majoreffortof thephasei study(ref.I) was to compile,edit,andpreparefor publicationas a NASAcontractorreportthe existing-informationon the B-I SMCSconceptualdesignand development.The majoreffortof thepresentphaseII studyis to reporton the analysesand flighttestsof theSMCS;existinginformationhas been augmentedby someadditionallimitedanalysesand flight-testdatareductions.
FLEXIBLE AIRCRAFT EQUATIONS OF MorION
The equationsof motionof the flexibleB-1 formthe foundationof muchthat is to followand so it is appropriateto discussthemfirst. The treat-ment of theseequationsis suchthat theyarenot developedhereinfromfirstprinciples.Thereare manytextbooksthatdo thistask;references2 and 3are typicalof these. The equations,however,are presentedin sufficientdetailto be readilyrelatedto the textbooktreatments.The equationsof
3
motion oi,this sectionof the reportwere developedto serve the purposesof ride quality,terrain-following,and handlingqualitiesevaluations;theequationsof motion relatedto the loads analysesare discussedin a subse-quent section. An attemptwas made to includethe main aircraft charac-ter(stiesimportantto these severaltypes of studies;however, as specificstudieswere conducted,minor modificationswere often made. It is not intended
to catalogall of these iterations;where importantto the studiesreportedherein,they will be discussed.
The discussionstouch upon the key featuresof these equationsalong withexplanationsof the form of the data where it is felt that this would be help-ful. The authorshave electedto stay away frommatrix notation in these _initialdiscussionsin order to show as much informationabout the dynamicmodeling as possible. The appendix containsa list of symbolsused in definingthe equationsof motion of this section.
The equationsof the motion in tables I, II, and III are written in abody axes system (figures2 and 3) where the X-axispasses throughthe centerof gravityand is parallelto the vehicle fuselagereferenceaxis (FRL). Tohelp those who are more familiarwith stabilityaxes notation,it shouldbeobservedthat all of the aerodynamiccoefficientsbut two appear the same ineitherthe stabilityaxes systemor the body axes system. The two that aredifferentare the normal force coefficient,CN (bodyaxes), versus lift coef-ficient,CL (stabilityaxes), and the chord force coefficient,CC (bodyaxes),and drag coefficient,CD (stabilityaxes). For small angles of attack,CNnearly equalsCL and CC nearly equals CD.
All accelerationsand velocitiesare defined positiveas indicatedinfigure 2. One exceptionis noted relativeto the definitionof the verticalload factor. As a concessionto stabilityand controland flight test conven-tion, a positiveload factor,nz, is associatedwith a positiveCN. In con-trast, the definitionof Z-axisacceleration(consistentwith the positivedefinitionsof figure 2) is associatedwith a positiveCZ in figure 3
(Cz = -CN).
Ride quality analyses have been conducted on the simulator using time-domain equations while frequency-domain analyses have been conducted usingdigitalcomputingequipment. The equationsin the time domain are presentedin table I and the frequencydomain in tables II and III.
The structuralflexibilityof the air vehicle is defined in terms offree vibrationmodes of the structure(oftenreferred to as normal modes).Figure 4 definesthe sign conventionsused in associationwith.these normal
TABLE I. - GENERAL FLEXIBLE VEHICLE EQUATIONS OF MOTION, TIME DO~~IN
[Total Vehicle, Body Axes, Units: ft, lb, rad, sec]
FORCE EQUATIONS, RIGID-BODY MODES
- ~~~&H - (rr:ov:cyXco5r)6cv- (:~h.)5A - (t~j[(CN~TF"tJ
i3 - -', + Poc -~~),8 + ~?:,cosesln<p +(~ s..,\{C~ (J + CY.6Ibw)~ + Cyr('b..,')r- + CYf(bw'f~v. V) \m"l."J (0() ~~ ~ z.\i.J'
m n
+ CYp(~)p + Cyp(~fp + [[CY71lni+ CYni(~il + [(CyPn.n~(3 + Cyo..6r- + C'ys..Sr--.-. .=. I •ANrlt.YMI1ET~IC SYMMETRIC
+ Cy~~ +CY6~S: + CY&:v0~v+CY6~~CV +e;,OAoh.+c,-~6A.} +(~,jj&' +(~~(s\nf')~v+ (m:'I..~)8A
+ (~~)~(yv,)v"J
TABLE I. - Continued
STRUCTURAL_DE EQUATIONS
Symmetric modes, i --1 to n
_._..___C_ _nJ=.l
\ MI \ MI
Antisy_etric modes, i = 1 to m
m n
+ _,_vJ__ + +J=l ,J=l
_T "Am 0 "
TABLE I. - Continued
b,_]B_TEQUATIONS, RIGID-BODY MODES _ Cm_(_--" + \ ly /L (o0 .
n
\- Iy
- +\_,(i,.
_w_.r_,Syl.lrl i_T RI C S'Y_11"t£TRt¢"
.i_ _. \ Ix / lx /
TABLE I. - Continued
LOAD-FACTOR EQUATIONS (ACCELERO_TER READINGS) _ .xX
ROMi_" 7_I AT ANy | r_ is-_-IN THE SAME"rATION )4..Y.i2! DII::_ECTION AS -I- Cr. I
TATION X,Y, ;z I'l), IS + IN THE SAME
DII_ECTION AS -i- C7ROTARY-RATE EQUATIONS (RATE GYRO READINGS)
n
[_,T ANyTATION _(,y] i=l
_--_L i ._,g_ .
P = P + Le, r_;('._._-,_._R,_.o0_s)rATANy X'Y_]_JLSTAT ION
i,,I
I-- = I_ + _-=-_Y'_ "rAT ANY Z__' 7_i (ANTISYT'II_ETRIC NODES)LSTATI ON _,Y,_] i=l
TABLEI. - Concluded
EULERANGLES
COS_
EARTH-AXESVELOCITIES
_-- N_CoseCOS_+ Vy(sIn_slnocos_- cos_sln_) +V_(co%_slnocos_r. sln_ s,n_)
V..E--_coses,n_. ',_(cos}cos_+s,nc}_nes,mf)+V_(cos}s,nesmg-sln{cos_r)
x4=-_s_ne+vrs,n_cose+v,cos}coseA convenient equivalent expression for Vze is in terms of altitude rate of change and _ and 8.
I_ = V. (s,necos0ccos/s -cosesln{s,rtp - cosecos_cos,_sln0c)
_o
TABLE II. - EQUATIONS OF MOTION FOR LONGITUDINAL RIGID-BODY AND¢_ Slqv_4ETRICSTRUCTURAL MODES, FREQUENCY DOMAIN
[TotalVehicle,Body Axes, Units: ft, Ib, rad, sec]
NORMAL-FORCE EQUATION, RIGID-BODY _K)DE
PITCHING-MOMENT EQUATION, RIGID-BODY MODES
TABLE II. Concluded
SYHMETRICSTRUCTURALMODE EQUATION,i = 1 to n
,<,.+i<...>t,_ I%7 t,i<,U"n H x#
,/NORMAL LOAD FACTOR (INCREMENTFROM 1 g TRIM) t_,_
FArA.,_x.¥,lz-J i-, iL_TAT IONNote: The variablesare complexin form but
have been written as _, q, _i, 6k, andwgas a space-savingnotation.
[( ) ()R + i ( ) _()g]representsreal andimaginaryparts of unsteady aerodynamicscoefficientform.
These equationsrepresentperturbationsfrom a 1 g trimmedflight condition.
• F_a
N TABLE III. - EQUATIONS OF MOTION FOR LATERAL-DIRECTIONAL, RIGID-BODY, AND
ANTISY_4_ETRICSTRUCRJRAL .MODES, FREQUF_NCYDOMAIN
[TotalVehicle,Body Axes, Units: ft, Ib, rad, sec]
SIDE-FORCE EQUATION, RIGID-BODYMODE
!
TABLE III. - Continued
ROLLING-MOMENT EQUATION, RIGID- BODY ~IODE
ANTISYMMETRIC STRUCTIJRAL MODE EQUATION, i = 1 to m
TABLE III. - Concluded
LATERAL LOAD FACTOR
STATATANy
_oN x,y, :_] i:l
Note: The variablesare complex in form but
have been writtenas _, r, p, qi, 8k,and Vg as a space-savingnotation.
+
[( ) ()R i ( ) ()i]representsreal
and imaginaryparts of unsteady aero-dynamicsin coefficientfom. Theseequationsrepresentperturbationsfrom1 g trimmedflightconditions.
+Y
+q
+vz,a
+X+r
+P
'+Z
_ Bqdy axes
Figure2. - Sign conventionfor rotationaland linearrates and accelerations.
+Cn +_I_
+N .6r+Y
+C N
+N .._ +C_- +L
' = Differental angle6H
+T between left and\ right panels
+Cz+Z
+X +6cv, Symmetric deflection +Z
+_cv, Differential deflection
Deflection magnitude and sense
referenced to right panel; right +Cm
panel trailing edge down, left +M
panel trailing edge up.
Body axes
Figure 3. - Sign conventionfor coefficientsand controlsurfacedeflections.
Symmetric Modes
Side view
+_fx
+ X Body axis_.d d,
ing point [ +Crli
AntisymmetricModes
+_/i_lizing point +C7/.,
<+x ' --_'x cG _+4 'xTop view
Looking forward
Figure 4. Structuralmode deflections,slopes,and generalizedforcessign convention.
17
modes. There are a number of advantagesto using the modal approach incontrastto the direct-influencecoefficientapproachfor such studiesasreferred to herein.
(i) It can describe the static as well as dynamic characteristicsofthe flexiblevehicle in a consistentmannerwithin the same format.
(2) Both inertiaand aerodynamicloadingsare accountedfor in themodal generalizedforce data for both the static and dynamiccases.
(3) The modal data are amenableto a number of simplificationschemes.
(4) The approachintegratesbest with control systemdesign requirementsin the handlingand ride qualitiesdesign areas.
Considerfirst, the most extensivelydetailed equationsof motion in thetime domain as describedin table I. These equationsneed be used in this
detailedform shown only when large-scalemaneuveringis studiedas in theterrain-followingproblem. More simplifiedequationswill serve other analysespurposes.
These equationswere developedassumingthat the angles of attack,_, andsideslipangles, 8, (figure5), would be small (lessthan i0 degrees),but thevehicleorientationin space as definedby the Euler angles9,@ , and • (fig-ure 6) would be unrestricted(exceptfor @ = 90 degrees).
Aerodynamicdata indicatedare preliminaryestimatesof those required.These data are shown in derivativeform exdeptwhere it is anticipatednon-linearcharacteristicswith _, _, or controldeflectionoccur. As an example,the normal force curve was expectedto be nonlinearwith _ so the normalforcecoefficientis expressedas CN(_) insteadof the more familiarlinear form
CN_-
The controlsurfacesexplicitlyshown are those anticipatedbeing requiredby either the Stabilityand ControlAugmentationSystem (SCAS)or SMCS. Other
control-surfaceinputsare shown in generalform as functionsof _k for thekth surface.
Becauseof the requirementto control structuralmotion at the frequenciesof the lower free-vibrationmodes, it is necessaryto considerunsteadyaero-
dynamic effectsof the controlmotion as well as the inertiareactionforcesof these surfaces. In the normal force-equationformatof _.ableI, the
unsteadyaerodynamicsare shown by the notion CN_k_k + CN. & for the kth6k k
18
E_
Body axes
Figure 5. - Angle-of-attack(_) and sideslip (8) definitions.
19
Vertical plane
Airplane plane
of symmetry
l
+ X-BodyI'II II Horizontal plane
m
I+ Y.Bod \ I
l \ I
\\\ + X-Earth
+ Z-Body+ YrEarth
I
3
+ Z-Earth
Rotation sequence indicated by numbers at arrow heads
Figure 6. Euler angle definitionsand rotationsequence.
2O
control surface. The control-surfaceinertiareactionforce in this equation
has the form (mk _k) 6"k- Similaraerodynamicand inertiareaction-forceterms may be recognizedin the moment and structuralmode equations.
Terms in table I involving4, 4, r, and p are generallyof small conse-quencebut have been includedto be consistentwith the frequency-domainequationsof tables II and III where unsteadyaerodynamicsmust be considered.The desirabilityof this consistencywill become apparentas these equations
" are discussedin subsequentparagraphs.
The B-I enginegyroscopicmoments, IR_Rr and IRO_Rq, shown in the equationshave not proved to be significantin handlingqualitiesor terrain-followingstudies. It had been anticipatedthat they might have been significantinlarge-scalemaneuvering. They have, however,been left in the equationsoftable I.
The equationsof table I includethe abilityto change speed. It isassumedthat if significantchangesin Mach numbersare to be realized,thesechangeswill be reflectedin the appropriateuse of Mach carpet data for theaerodynamicsrequired. Velocity changeswill show up directly in Vo, while
combinedaltitude-velocitychangeswill appear indirectlyin qo, (1/2 pV2).
The gust representationshown in table I is in the aerodynamictransfer-
functionform. Tilegust excitation,Wg and Vg, would come from random signalgenerationsources,shaped and scaledto reflectthe desiredgust powerspectraldensity and intensity.
The longitudinaland lateral-directionalrigid-bodymotionsare coupledduring large-scalemaneuversthroughthe inertiaterms and enginegyroscopiceffects. The symmetricand antisymnetricstructuralmode motionsare coupledby terms representingthe dihedraleffectdue to symmetricwing bending. Forsmall motions about a trim condition,the rigid-bodymode equationscan bedecoupledthrougheliminationof the inertiaand gyroscopiccoupling. Thestructuralmode equationscan be decoupledby using trimmedairplane static
-: symmetricstructuralresponseparametersat fixed values to determinetheeffectivedihedraldue to symmetricwing bending.
The load factor and rotationrates of the large-scalemaneuveringflexibleair vehicle (as read by accelerometersand gyros mounted on the fuselagestructure)are presentedin table I.
Euler angle equationsand earth axis velocitiesare given in table Iand can be used in terrain-followingstudiesto determinethe vehicle attitudeand locationwith respectto the earth'ssurface.
21
The frequency-domainequationsof motion for the flexibleairplanearegiven in table II for the longitudinal-sy_netriccase and table III for thelateral-directional-antisymmetriccase. These equationsare uncoupledandrepresentmotion perturbationsfromwing-level1 g trimmed flight. Compari-son of these frequency-domainequationswith the uncoupledtime-domainequationswill help in identifyingequivalentterms.
As shown, the vehicle responseaerodyn.amics(that is derivativesassociatedwith responseparameterssuch as a, _, q, q) are quasisteady,while the controlsurfacesand the gust are sho_cnas functionsof the forcingfrequencyin theform [( )R + i ( )I]- This formathas been convenientand sufficientlyaccurate for.preliminaryride qualitiesand structural-modecontrolanalyses.Digital programsare availableat Rochcell;however,that will also acceptvehicle responseaerodynamicsfrom unsteadyaerodynamictheoriesas a functionof frequency,permittingmore refinedride qualityand S_ICSstabilityanalyses.
The equationdescribingthe normal load-factorresponseat any locationin the flexibleaircraft is presentedin table II. The similarequationforthe lateralload factor is given in table III. These responsesare used inride qualityanalyses.
FLEXIBLE AIRCRAFT ANALYSES bDDEL
This sectiondescribeshow the data were obtainedto implementthe flexibleaircraftequationsof motion used for ride quality analyses. It will not bethe purposeof this presentationto providea complete set of data used inall analysesdiscussed;but it will be the intent to provideunderstandingofthe data used.
DYNAMICANALYSISSYST_
As a basis of understandingthe contentsof this sectionbetter,as wellas topics of other sections,the chart of figure 7 is presented. Shown is --the completedynamicanalysis system supportingthe developmentof flexiblevehicle dynamicanalysismodels for controlsystem developmentand ride qualityanalysesat RocMcell International'sNorth AmericanAircraftDivision,E1 Segundo. The path throughthis system as employedin developingthe SMCS isas follows. (No supersonicanalyseswere conductedduring SMCS developmentsothe Mach-box programcapabilitywas not used.)
Starting at the top of the chart, it is seen that the processbeginswith a definitionof the vehicle geometry,basic wind tunnel correlatedaerodynamics,structuralstiffness,and mass characteristicsbeing providedto
22
I I I il
Vehicle geometry I- Basic aerodynamics
Structural definition
I-- J T IMaSs characteristics,_
'' _ ' ,
STAR 6 stiffness'
vibration analysis El, GJ'SIC
Modes, frequencies
J Modal Iprocessing
_ Normalized modes,
Doubletlattice II lJ M>IMaChbox l_lizedm<] l I
Dynamics,J _ ' I_ , Generalized
Technolog_ J Generalized j aero forces
I I I
Control _ FH251/FH255 H STAR 4
system Dynamic Flutter
analysis response analysis
Control system requirements Flutterdynamic response, stability characteristics
• _ I al
-, Figure 7. - Dynamicanalysis system.
23
the DynamicsTechnology(,roup.As indicated,the stiffnesscan he used inthe EI-iU formator in the formof structuralinfluencecoefficients(SIC).The ,lassand stiffness&ata enter the STAR 6 program,and vibrationanalysesof the whole vehicleare accomplished. The output of the program,then, iswhole vehiclevibrationmode shapesand frequencies.
These mode-shapedata are next manipulatedto producenormalizedmodes.For tileB-I ride qualityanalyses,both the sy_netricand antisymmetricmodeswere normalizedto a I_int on the nose of the vehicle. Generalizedmass dataconsistentwith the normalizedmodes are produced. Finally_modal deflectiondata are developedthrough interpolationprogramsalong selectedstreamwisestrips for input into the aerodynamicprograms.
For tile ride quality related studies reported herein, the Doublet Lattice.Program has been used to provide theoretical aerodynamic generalized forcesas required. These generaI ized aerodynamic forces are in dimensional form.Programs have been developed which process the data from dimensional form tothe coefficient form required by the dynamic analyses programs. These dynamicanalyses programs employ the frequency-domain equations of motion discussedearl ier.
The 1:tt-251 program provides dynamic response results for the longitudinal-symmetric case, while the FH-255 program provides dynamic response results forthe lateral-directional-antisymmetric case. Both of these programs can accepteither frcxtuency-dependent or quasisteady data. Active controls can beincluded. Frequency responses due to gust or control forces may be obtained.When gust inputs are employed, ride quality parameters are output and control-system deflections and rate responses are obtained in power spectral densityform. Stability analyses are performed using the characteristic determinantfrequency evaluation technique of reference 4.
FREE-FREE VIBRATION MODAL DATA
The flexible aspects of the aircraft have been treated in the modal _format as opposed to the direct-influence-coefficient approach. Ninety per-cent of the ride quality analyses performed have been accomplished usingfree-free vibration modes which were obtained using an EI-GJ description ofthe vehicle stiffness; more recent modal data have been obtained usingstructural influence coefficients. The details of the EI-GJ approach arediscussed here.
Figure 8 shows the typicaldistributionof mass pointson the elasticaxes assumed. This is an earlymodel; figure 9 shows a refinementof thefuselage elasticaxis made at a later date. Each analysisincludeda flexible
24
. _FS 2461 (969) FS 4018 (1582)
Nace IIe
, points
Z
Typical control point motions calculated from EI-GJ approach
WL 72.4
IUk -81 (-32) FS 3942 (1552)
Figure 8. - Typicalelasticaxes and mass point locations.
25
I-o
Elastic axis., 'D f EA _ FRP
I -- L._ _1 (w,o)
Elastic axis coordinates*
Point Fus Sta WL
A O 38.6 (15.2)
B 406 (160) 34.0 (13.4)
C I003 (395) 143.8 (56.6)
D 1283 (505) 123.2 (48.5)
E 2477 (975) 80.0 (31.5)
F 3937 (1550) 154.9 (61.0)
G 4013 (1580) 50.8 (20.0)
H 4430 (1744) 57.7 (22.7)
*Straight lines between these points
Figure 9. - Fuselageelasticaxis refinement.
wing, fuseIage, horizontal tail, and vertical tail. Also included wereflexibly mounted engines/nacelles. The resuItant mode shapes consisted ofelastic axis deflections and rotations illustrated in figure 8.
Because the analyses were primarily oriented toward ride quality at thecrew station, the free-free vibration modes were normalized at the most forwardmass point at the nose of the aircraft. While most modes show a high degreeof coupling among vehicle components, table IV identifies the main component
; (wherethis is possible)and lists the associatedfreque.ncies.In the analysesdiscussedherein, I0 symmetricmodes and 12 antisy_netricmodes were used.Discussionsto follow later in this sectiondescribethe rationaleused in
selectingthe modes shown.
When groundvibrationtest (GVT)data became available,they were usodto upgrade the modal data. Symmetricorthogonalmodes were successfullyex-tractedfrom such tests. First, the effectsof the soft supportsystemwereremovedfrom the measuredorthogonalset of modes. These data were next used,with proper fuel weights included,to analyticallyobtainedorthogonalfree-free modes at the desiredweight condition. Itwas not possible,however,to directlyextractan orthogonalset of antisymmetricmodes from GVF data.To obtain usable, consistent, antisymnetric modal data reflecting test results,a technique was used of adjusting local stiffness data until a successfulapproximation oF the measured data were obtained analytically. Then, as inthe s_nmetric case, the effects of the soft suspension were deleted and thedesired fuel weights adeed analytically to obtain orthogonal antisymmetricmodes. The data of table IV are typical of those based on the GVT data ob-tained as described. It is to be noted also that structural damping wasextracted; these data were obtained using the oscillation decay method.Figuresi0 and II present typicalsymmetricand antisymmetricmode vector plots.
The computerprogramsused to determinethe aerodynamicdata requirevibrationmode shapes to be definedon a grid systemwhose chords are parallelto the free stream. Thus the basic mode shapeswere interpolatedto find
point deflectionsalong strean_isechords for all liftingsurfaces. (See fig-: ure 12.) The mode-deflection data on the grid system shown were used as input
to the Doublet Lattice Program where interpolations to the Doublet Lattice gridsystem were made and required slope data determined within the program.
AERODYNAqICDATA
The DoubletLatticeaerodynamictheorywas used to obtainmost of the
aerodynamicsdue to the flexiblestructure. In addition,extensivewind
27
TABI,Ii IV. - TYPICAl,ANALYTICALSTRUCTURAl,MODEI)ATA
Wt = 119,296.8 kg _(263 000 lb)A = 65°
Structural ablode Mode Frequencyno. description Hz damping
gs
Symmetric7
1 Wing, first bending 2.22 0.0622 Fuselage,first bending 2.84 .0943 Fi'rstnacelle 3.26 .024
4 IIorizontal tail, f_rst bending 4.19 .0285 Wing, fore and aft bending 4.23 .0526 Fuselage, second bending 6.28 .0167 Wing, second bending 7.57 .0228 llorizontal tail, fore and aft bending 8.31 .0649 Fuselage, third bending 11.15 .055
10 llorizontal tail, first torsion 27.35 .042
Antisymmetric
1 First nacelle I.73 .145
2 Wing, first bending 2.41 .0543 Horiz tail, first bending 3.51 .0434 llorizontaltail, first bending 3.96 .0255 tlorizontal tail, fore and aft bending 4.14 .0496 Wing, fore and aft bending 4.20 .0317 Fuselage, first bending 5.58 .0328 Wing, second bending 6.96 .0319 Vertical tail, first bending 7.21 .019
10 Fuselage, second torsion 9.72 .07811 Fuselage, second bending 10.30 .02212 Vertical tail, first torsion , 35.34 .020
aDeterminedfrom groundvibrationtests
28
Fore and aft motion
Lateral motion
Vertical motion
Mode l
_ Frequency = 2.23Hz
Figure I0. Typicalsymmetricmode vector plot.
29
Fore and aft motion
Lateral motion
Vertical motion •
\
\
\
_/ Mode l
Frequency = 1.73 Hz
t :-
Figure II. - Typicalantisymmetricmode vector plot.
3O
..
9
17
;>..1
' ..'..
t. 3Sb
03 ••
Figure 12. - Typical structural mode deflection grid points.
Horizontal tail location,
WL 320 (126)
WL 63.5 (25)
I FS 0
Horizontal tail pivot,
FS 4018 (1582)
Figure 13. - Panellingand box grid for Doublet Latticeaerodynamicsprogram.
t_mel data were also generatedto obtain static,rigid-bodyforce and moment-coefficientdata along with pressure-distributiondata. These two data sourceswere correlated (to be described)and used as input to the flexible aircraftanalyses.
The panelling,box grid, and control-surfacessetup for the DoubletLatticeProgramare shown in figure 13. The wing and fuselageforebodyhadfivemain panels with a total of 151 boxes. The horizontaltail had twopanels and a total of 64 boxes; the horizontaltail was _ut in, also, as anall-movablecontrolsurface. The vertical tail had two panels and 73 boxes.In addition,the vertical tailhad an end plate at its base consistingofone panel with 30 boxes. As shown in figure 13, the lower rudder controlsurfacewas also modeled.
For longitudinal-symmetricdata, includingcontrol effectivenessand gustdata, the wing/forebodyand horizontaltail were run as shown in figure13.This same wing/forebodyarrangementwas used along with the full empennage(horizontaltail, verticaltail, and end plate) to obtain lateral-directional-antisymmetric,whole-vehicledata. Vertical tail gust data, lower ruddercontroleffectiveness,and differentialhorizontaltail effectivenessdatawere obtainedby runningthe empennageplus end plate as an entity. Forfuselageside gust and generalizedaerodynamicforces,a modified slenderbody theorywas used which made use of wind tunneldevelopedside-forcedistributiondata.
It is to be noted that the SMCS vanes were not modeled for the Doublet
LatticeProgram. The aerodynamicsfor the vane were obtained fromwind tunneltests as describedin referencei. None of the vane-relatedaerodynamicswerefrequencydependent. The reducedfrequency(k) determinedon the basis of anassumedfrequencyof i0 Hz (62.8 rad/sec),the highestfrequencyrange thatvane is expectedto be effectiveat M = 0.85, is
k = _--_= (02.8)(2.46) = 0.0812: ZVo (2](9Sl)
Compared to k for thewing under the same circumstances
k- (62.8)(15.23) _- 0.503(2)(9Sl)
This is a relativelylow reducedfrequency,and the aerodynamicunsteadinesseffectsare judged to be acceptablysmall. In retrospect,this appearstohave been a valid judgement. In generatingthe structural-mddegeneralizedforce coefficients,the vane force was assuned to act at a point; i.e.,
33
FS 581.66 (229). Thus C_. = ¢iFS 581.66for symmetricmodes is anexample. ' 1_cv CN_cv
In developingthe DoubletLatticepanelingand box patterns,lifting-surfacespanloaddistributionswere comparedto the wind tunnel-relateddata.The gap betweenthe vehiclecenterlineand the first row of chordwiseboxesof the horizontaltail is one of the devicesused to obtain matcheddata sets
for the horizontaltail. It was reasonedthat with good matches of thelifting-surfacespanloadings,the computermodel developedwould give validanswers for both rigid body and structuralmodes.
The point of vi_¢ was adoptedthat the wind tunnel-relatedairloaddis-tributionfor the rigid vehiclewere the most accuratedata available. Thus,
the rigid vehicleaerodynamicscouplingsinto the structuralmodes (Cni_ and
Cni6 data are examples)were computedusing these distributionsratherthanDoublet Latticetheory,for the zero-frequencycase. These data, in turn,
were used to scale the frequency-dependentdata produced by the Doublet
Lattice Program. The Cnin.jand Cni_j data were used directlyas generatedbythe programwithout scalingfor both the synraetricand antisy_netricmodes.
As an exampleof how the wind tunnel data and the DoubletLatticefrequency-dependentgust aerodynamicdata were brought intoagreement,considerthetypical e_ampleof figure14. Shown is the pitching-momentcoefficientdatadue to a unit vertical gust velocity. The magnitudetrendsversus the reducedfrequency,k, were assumedbasicallycorrectand all adjustmentsto match windtunnel based data were made at low frequenciesas illustrated. The data magni-t_le of the real componentat zero frequencywas adjustedto match the windtunnel based data and then faired into the basic real curve at low reduced
frequencies. For the case illustrated,the gust coefficientat zero frequency
was determinedfrom angle-of-attackdata, CmWg (Cm_)/Vo. This matchingofthe angle-of-attackand gust velocitydata is essentialto obtainingvalidpower spectraldensityresponsesdue to gust data at low frequencies;a mis-match will producea load factorresponseat zero frequencywhich is not therein the real world. In the realworld, a stable aircraftwill weathervaneintothe resultantvelocitydue to the combinedmotion and gust velocity and haveno load factorat zero frequency.
As mentionedwhile discussingthe analysesflow, the data producedbythe DoubletLattice Programmust be reducedto the coefficientform of theequationsof motion discussedin the previous section. In order to do this,the data are processedin the followingmanner.
34
0 Real l Symbols denote
[] Image[ calculated data pointO.OO4
0.002
Crr_g
1 ,
ft/sec 0 __ .6 k .8 1.0 "_ I._I 6-0.002 ,--_r]F -- ..
-- ,ow-freqoencyadjustmenttofit wind-tunnel-test based data
-0.004
Gust reference point is FS 2649 (I043)
Figure 14. - Typical low-frequencyadjustmentmade to analyticalfrequency-dependentaerod_lamicdata.
The programoutputsdata normalizedto air density (2p)and to frequency
(o_2) as illustratedhere using the dimensionalliftforce due to plungingLh(forthe whole vehicle), [Lh/2P_]. The divisionby two is because the data
generatedin the program are for half of a vehicle, p is assumedto be unityby the program. Velocity is determinedby the data of Machnumber and velocityof sound at the altitudeselected. Frequency,o_,is input at a number ofselectedvalues.
Having this information,table V illustrateshow the frequency-dependentcoefficientdata are developedas a functionof frequency(mr reducedfrequency,k). Longitudinal-synmetricalcoefficientsare obtainedusing the informationof table V; lateral-directional-antisymmetriccoefficientdata are obtained ina similarmanner.
Reflectedin table V is the sign conventionbuilt into the Doublet LatticeProgramat Rockwelland the sign conventionassumed for the equationsofmotion of tables I, II, and III. The or_l°ydifferenceof importanceto theunderstandingof the derivationsof table V is the fact that lift force,L,of the DoubletLatticeProgram is of opposite sense to the equationof motionnormal force, N, and its associatedcoefficient,CN. Otherwise,the pitching
moment, M, and structuralmode generalizedforces,Qi' are identicalindefinition.
As an aid in understandinghow table V was assembled,the following
exampleis given for the derivationof CN.
For no pitching (@ =@= 0)
Vo Votherefore:
qoSw CN _= - [Lh] h force unitsIMAG
IMAG
v°2T CN= -o!
IMAG
INAG
36
TABLE V. LONGITUDINAL-SYMVIETRICAERO COEFFICIENTS FROMDOUBLET LATTICE PROGRAM
37
The bracketedquantity [ ] here and in table V comes directly from the DoubletLattice Programas a functionof frequency.
One of the questions to be resolved in using a modal approach to aircraftflexibility modeling is the one of how many modes to use. The technique usedto help make this decision for the described analyses is discussed here.First, the following criteria were developed as guides in the decision process.The modes contributing to main a_roelastic impacts on all short-period andDutch-roll characteristics were to be included. The modes contributing tomain aeroelastic impacts on control effectiveness were to be included. Themodes contributing most to flexible fuselage motion at the pilot station andactive control sensors were to be included.
Figure15 is typicalof the data generatedto assist this mode selectionprocess. For the longitudinal-syn_netriccase used as an illustration,some
25 whole-aircraftnormalizedmodes were input to the Doublet LatticeProgramand run at a frequencyof n/2 (a frequencyjudged to be in the regionof short
period and Dutch roll frequency). These data then were reducedto aerodynamiccoefficientform and enteredinto a programwhich calculatesquasisteadyflexible-to-rigid(F/R) ratios (referto 'TruncatedAnalyticalModels") for
all of the key aerodynamicderivatives(CNaandCn_ are used as illustrationshere). These F/R ratios are calculatedas one mode after the other is elimi-
nated. As a result,it is possibleto identifythe individualmodes contribut-ing most to a given derivative'saeroelasticimpact.
After the preceding procedure is accomplished for all important deriva-tives, plots similar to figure 15 are assembled and inspected as a whole. Thus,those modes making important contributions to all derivatives are selected forr et ent ion.
As far as fuselagemotion was concerned,as many modes as possible,havingfuselagemotion as a main component,were selected. Usuallymodes reflectingup to the third fuselage-bendingand the second fuselage-torsionmode couldbe selected.
• 2-
CONTROL-SURFACES INERTIA REACTION FORCES
The inertiareactionforces of the controlsurfacesare importantinputsto the stabilityanalysesof active controlsystems. The detailedfinal form
of these inputs for the B-I controlsurfacesare indicatedin the equationsofmotion of the first part of this report. The basic approachthat was used indevelopingthese expressionswill be discussedhere, but each control-surfaceinput will not be developedin detail.
38
For constant Mach number and altitude
l.O
0.8
CNol_ _-- ....F/R %,
%..-_
0.6 \\llll
0.4 |l -. /iI /"L_....----/ CM_x
0.2
0 I * I | | I I a s i I I I I I I I I I I I I I a
0 4 8 12 16 20 24
Number of modes
"L
Figure 15. Typicalaeroelasticflexible-to-rigidratiodata for aerodynamic
coefficientsas a functionof participatingstructuralmodes.
39
Figure 16a shows that when a typicalcontrol surfaceis acceleratedinthe positivesense of the deflection,a mass reaction force and moment aredevelopedat the centerof the mass of the controlsurface. This force andmoment are reactedinto the basic aircraftstructureat the control surface
hingelineas a force and moment as shown. The force acting on the aircraft isthe force shown,andthe moment acting about the aircraftcenter of gravityis (_m_+ I )6" where _ is the distancebetweenthe control-surfacehingeline_Eand the aircraftcenter of gravity.
The generalizedforce acting on a typicalstructuralmo'dedue to control"surfaceaccelerationis illustratedin figure16b where the specific exampleof a s_etric mode is used. Before getting i1_tothe specificexample,con-sider the conceptof a generalizedforce. A generalizedforce has units ofwork, that is m - N (foot-pounds). In this case, it is eitherforce multipliedby mode deflectionat the force applicationpoint,moment multipliedby modeslope at the moment applicationpoint, or both. A positivegeneralizedforcewould act to increasethe deflection (or virtualwork) of the structuralmode.Lookingat the example in the figure,it can be seen, for the example shown,that the reactionforce causesa generalizedmode force incrementof -_i_Lm_"
and the reactionmoment a generalizedmode force incrementof +¢iILI_L6".
The precedingdevelopmenthas proved an adequaterepresentationfor smallcontrol surfacessuch as the B-I SMCS controlvane and lower rudder control
surface. It is not accurate enough,however,for large control surfacessuchas the all-movablehorizontaltail. In this latter case, it was necessarytobreak up the mass characteristicsof the surfaceonto a distributedgrid system.Using the distributed:m_ssdata and the previouslydevelopedlogic, generalizedforceswere developedfor rigid body and structuralmodes. These data, devel-oped using distributedmasses,were input to the digitalprogramwhich imple-ments the equationsof motion of table II using an equivalentpoint mas.srepresentationof the data.
During the B-I development,checksof the pitch SCAS, with the aircrafton the ground resting on its landinggear, revealed a structuralmode-couplinginstabilitywhen excitedwith sharp horizontaltail Controlinputs. This • _-instabilitycould be analyticallyduplicatedby using generalizedcontrol-surfaceinertiareactionforces developedemployingthe distributedmassapproach but could not be duplicatedusing the single-pointmass representation.
The inertiareaction forcesare particularlyimportantto the stabilityof SMCS with the aircrafton the ground. The inertiareactionforces have anopposite s_se to theaerodynamic forces of the SMCS vanes. The stabilityofthe system is establishedby these aerodynamicforces. Thus, if the aerodynamicforces disappear,the feedbacksense is effectivelyr_versed,producingan
t
4O
Reaction force and moment actingon primary structure at H_
Reaction force and moment
about surface CG
° _ Primary
structu
I ' +(_
'IAircraft H_
, [ _l_._.._l CG of control
l l surface mass,m
Force at aircraft CG = +m{_"
Moment about aircraft CG = -(_m{ + IFL )i_"(pitch axis assumed
for illustration)
(a) Rigid-body generalized forces
Symmetric mode i
\ Reaction View from
X 4///_/_'(t s'CG Reaction"force _iFL left sidemoment \_
7L
Z
Note definitions of +(_liand +_iI'L.
q, 'm //5. Structural generalized forces =-( { _i H_
(b)Structural generalized forces
Figure 16..- Typical controlsurfaceinertiareactiongeneralizedforces.
41
instability. On the B-I, a switchon the landinggear preventsoperatingtheSMCS while on the ground, precludingany inadvertantdamage due to thispotential instability.
ACTIVE GONTROLSYSTt_/IS ..
Two types of active control systemswere includedin the analysesper-tainingto this study. One type, SCAS, is associatedwith controlof whole-
vehicle (shortperiodand Dutch roll) modes of motion. The second type,SMCS, has the functionto controlfuselagestructuralmotion to improveridequality.
The block diagramsand analyticalmodelingof the SCAS are given infigures17 through19 and S_ES in figures20 and 21. Flightcondition-dependentgains are shown for M = 0.85 at SL. These figures indicatethe type ofsensors,compensations,gains, and actuatormodeling assumedfor each of theindicatedsystems. The control-surfacedeflectionequationsare cast in aform directlyusable by the Rockwellresponseanalysesprograms. That is tosay, the overallgain is indicated,system dynamicsare representedby numeratorand denominatorroots of polynomials,and vehicle motionsare definedasmeasuredby the appropriatesensors.
COMPARISONSOF ANALYSESAND FLIGHT-TESTRESULTS
The ride quality indices,_z and Hy (referto reference5 for detaildefinitions),for the vertical and lateralaxes, respectively,cannotbemeasureddirectly in flight. One difficultyis the fact that the parametersare obtained from weighted power spectraldensitycurves of crew-stationaccelerations,and the weightingcan presentlyonly be done as a postflightoperation. Secondly,it is very difficultto preciselyfly a specificationvehicleweight at the specificationMach number and altitude. For the B-l,the approach has been to demonstratethat the analyticalmodel can duplicateflight-testresultsand then proceedto use the verifiedmodel in the required - _ride qualityanalyses. It is the intent of this task to presentthe datamatches which providedthe verificationand give an evaluationof the factorsaffectingthe matches.
The flight-testdata obtained for matchingpurposeswere frequencyresponsesof load factorsat FS 571.5(225)due to SMCS vane inputs. Bothvertical and lateralload-factordatawere obtained. The flight conditionflown was M = 0.85 at 762 meters (2500 feet). The vehicleweight was approxi-mately 119 296 kilograms (263000 pounds). Wing sweep was 65 degrees. Datawere taken with all controlsystems inoperative;then, with only the SCAS
42
•8 p.Ao
R,Ki"E aG,_o _<p (s_.o4S+9co_ .a(, 5'z.14. 1o
k_%, I'2.1.,,,L"_ 4,9 _01 AT _+ 57,14- S_|OF.S. 2;<.49 5.k.
:q Hi NOTCH MPEMSATIO SEP-.VO ACTUA_W|LTFJ_
RAD
I NOR.HAL L__ GAIN / //3--__--'2-('Y_;%'7"°_6_r'i'13"28__.S.2'¢49
mAD 1199<3.;_3
"L_ ;=1 ..J
For flight condition of M = 0.85 at SL
Figure 17. - Pitch axis SCAS analyticalmodel.
~~ Fl XEO F'OR ALL
g J=LIGHT CONDITIONS~ .....~ .44- 6.0
YA\N~TE COMPEN~TIOH K~...G'fRO 57.14 20
~.s. :l.(.!l19 S'+ 57.14- 9'+20( (043)
I<h NOTCH SERVO30 P ~ILTER
OR.Sr30 1.0
ATRAn
LATERAl COMPfNSATION l<MI"l S.L..Acr:.EL. "
For flight condition of M = 0.85 at 5L
Figure 18. - Yaw axis SeAS analytical model.
_S+ 1.05±J20.97) R_oRAo/_SC 57.14 I0(,5+ %.4 + J 19.35) .5? S -i-57.14- _ + IO
F_ATE - G_|N SERVO TUATOI_
GYI_,o r F ILT E I_
E s._649 FIXED gO_. RAD(IO4.3) ALL FLIGHT
CONOITIONS
325.7 m
' (S+l.os+_2o.O_I ! r57 5714 I0 - •-... ,. <a": -( )( s+_.4-JLg.3_S_7._X_-i_/_,Co)_+(,.olP :__AD J= I
For flight condition of M = 0.85 at SL
Figure 19. Roll axis SCAS analytical model.
4_OX
_0o) los" _o-7.31 c_+to .5"¢.5' .159 _+.lo _.SO
- ,0d@FH I_ F i+AT COFt P. _ ILTF_.R ASHOUT Tt._TOW_HF_
F.S. 57|, 5 CONSTANT C.OC.KPIT,, S.L. HA_ sE'r'r_,_- DF,J3
IO (C_'H57_'1"i57) H=.8S" o1=,O=1 VOLT-cj+jo
,_+!! /NGI_III_I_ LAG I-"
c-e-I I 1I:'.s._
(_o,_3) _o.2.7
IS.7±iIS(..Z) S' I
I--'I
IOZ.7, I"I
, .(£ __I.('.Is3X_XI31XI_Io)OoXIoXso)_I V, \(_+l£7+-JISt.._)/______/I IXI_. _,_<.__,_,/'_IO43 _. "L
RAD i=l
For flight condition of H = 0.85 at SL
Figure 20. - VerticalSMCS analyticalmodel.
_8 los 5o-7.31 S'+ 1.0 .5"(.,,5" .153 _ + |0 $+50
AT _,_ V COMP. _,L'rE_. uA'_oII I I'_"_1I !1 I_
F..s.S'/!,s" ¢OHs'rAHT _.iS.7:!:_ 1_(..2) AT ¢<:x:KPrrIo (_41sTX_,+_s'f)M=.ss o_,o=IVo-T
_.+IO
I_IOHINALI-_ IL/_4_ _-- -- Note: Apparent extra -_ due to
c.G.I I I simulatio,ofbankanglein lateral load factorF._.Z_ terms by
Oo4s) s8.0_
_:_v/_- (S_.sX3z.z) .... S,I_(.S+IST_+ISZ)_+IO_+SoAS/ ° "_' & --'_ V.dV, i
eo._
• _I
RADFor flight condition of M = 0.85 at SL
Figure 21. - LateralSMCS analyticalmodel.
operating;and finally,with both SCAS and SMCS (cockpitgains 1.5) operating.The forcingamplitudeand frequencysettingswere set manually in the cockpit.The measureddata were processedto obtain both magnitudeand phase charac-teristics.
The analyticalstructural-modecharacteristicsfor the data-match
analyseswere obtainedfrom test and analysis sources. The symmetricstructuralmodes were obtained directlyfrom GVT. It was not possibleto obtain a set of
orthogonalmodes for the antisymmetriccase directly from the GVT. Instead,the analyticalstructuralmodel was adjusted in stiffnessuntil the frequencyand mode-shapecharacteristicswere as close to the observed characteristicsas
possible. In supportof these tests, the basic vehicleweight characteristics(no fuel) wePe identified. For the specificdata-matchinganalyses,fuelloadingswere determinedfrom measurementsmade duringthe flight-testperiodwhen the frequencyresponseswere being executed.
The pitch SCAS characteristicsused in the analyseswere as describedinfigure 17; the yaw and roll SCAS were as describedin figures18 and 19,respectively. The verticalSMCS descriptionis given in figure 20; and thelateral SMCS descriptionis shown in figure 21.
Comparisonof the analyticaland flight-testresultsof the normal loadfactor of FS 571.5 (225) _'requencyresponsedue to synmetricSMCS vane deflec-tions (SMCSused as an excitationsystem)for the conditionof no activecon-
trols (basicaircraft)are sho_cnin figure 22. The three responsepeaks are(startingwith the lowest frequency)first fuselagebending, second fuselagebending, and third fuselagebending, respectively;there is some influenceonthe second peak from the wing secondbending. The qualityof the match isconsideredexcellent. To obtain this match, however, three adjustmentsweremade to the originalmodal characteristics: (I) the wing first bending-modefrequencywas reducedfrom 7.57 to 7.00 Hz, (2) the fuselagethird-bendingfrequencywas reducedfrom 11.15 Hz to 8.60 Hz (thislatter change is substan-tial, and no reasonhas been found to explainwhy the originalmode was off),and (3) the structuraldamping (gs)for the fuselagethird bendingwas changedfrom 0.055 to 0.025. Havingmade these adjustmentsfor the basic aircraft .responses,no additionaladjustmentswere made to the analyseswith controlsystemsoperating.
Figure 23 displays the data matches for the frequencyresponsewith theSCAS operating. Comparingthe first peak-responsemagnitudeof this figurewith the previous figure,it can be seen that the SCAS excitesthis peak some.Again, the analytical-to-testdata match is excellent.
Figure 24 shows the data matches for the frequencyresponsewith boththe SCAS and SMCS operating. The effectivenessof the SMCS in damping thefirst fuselage-bendingmode is demonstratedby these data. The data match is
48
Note: 8cv is control surface deflection. The flight-test data measurementsof the forcing command were analytically processed to remove effects
of actuator dynamics,which were measured,in order to permit com-
parisons with analytical results on this and similar subsequentfigures.
150
" !oo /_/ \ I
// \Phase
ang]e 150 I \ , _\deg /j/ _/'
Sl I I I I I I _// _
O_ / 2 4 6 8 _'_0% !%I Frequency (f) - Hz
-50:-"
Flight-test data
Analytical data
8
grad
L
00 2 4 6 8 I0
Frequency (f) - Hz
M = 0.85, air = 762 m (2500 ft)Wt = 119 296kg (263 000 lb), .A= 65 °
Figure 22. - Comparisonof flighttest and analyticaldata, frequencyresponseof normal load factor at FS 571.5 (225)due
to SMCS vane deflection,SCAS off, SMCS off.
49
olOO !Phase I
angle ,50 /_/ ,__l_ 2//deg I0 / " a
_,/ 2 4 6 8 1
Frequency(f) Hz-50
Flight-test data
Analytical data
12 -
g
4
,d
0 I l I I I n n
0 2 4 6 8 lO _Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = 119 296kg (263 000 lb), A= 65°
Figure 23. - Comparisonof flight test and analyticaldata, frequencyresponseof normal load factorat FS 571.5 (225)due
to SMCS vane deflection,SCAS on, SMCS off.
5O
15o
Flight-testdata )SMCS vert gain 1.5
Analytical data 912
n z
8
g
4 / \\
7 "*
0 '_'I" I I I I I , , a i
0 2 4 6 8 I0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500ft)Wt = ll9 296kg (263 000 Ib), .A= 65°
Figure 24. - Comparisonof flight test and analyticaldata, frequencyresponseof normal load factor at FS 571.5 (225) due
to SMCS vane deflection,SCAS on, SMCS on.
• 51
excellentfor the first two responsepeaks and fair for the third peak. Thelack of better fit for this third peak was initiallyattributedto SMCS act-uator modeling in thishigher frequencyrange; however,using the flight-test-derived actuatormodel of figure 25 did not substantiallyimprovethe matchto the flight-testdata in the 8 to I0 Hz frequencyrange as shown in figure 26.
The comparisonof the frequencyresponselateralload factorat FS 571.5(225)due to differentialSMCS vane deflectionsfor the basic vehicle is shownin figure 27. The comparisonis fair; the frequenciesof the three peakresponsesare duplicatedby the analyses,but the amplitudeof the low-frequencypeak is off by a considerableamount. The phase angle is matchedreasonablywell in the midfrequencyrange only; specificreasons for why the data do notmatch better are not known. The peaks are identified(startingwith the low-frequencypeak) as the wing fore and aft mode with a large fuselageside-bendingcomponent,fuselagefirst side bending,and second fuselagetorsion. This lastpeak in the analysiswas obtainedby droppingthe frequencyfor this modefrom 9.72 to 7 Hz; although this mode was the only logicalone to adjust,there is no reason known for the noted discrepancy. Having made this logicaladjustmentfor the basic vehicle description,no other adjustmentswere madewhen the controlsystemswere operated.
The data comparisonof figure 28 are for the case of SCAS operating.The data are similarto the basic aircraftresponseof figure 27. Again,the agre_nentof analysisto test data is only fair.
In figure 29, the frequencyresponsecomparisonsare made for the caseof SCAS and SMCS operating. The trend of the analysesand the flight-testdata are similar;i.e., the first and secondpeaks are attenuatedbut thepeak around 7 Hz is increasedby the SMCS operation.
The implicationsof these analytical/flight-testdata comparisonsareimportantto the B-I ride qualityverification. The data imply that thevehicleanalyticalstiffnessand mass characteristics,whole-vehiclecontrol-surfaceaerodynamics,SMCS vane aerodynamicsand inertiareactionforces,andSCAS and SMCS modeling are fairlyaccurate. Thus, the ride qualitycharac-teristicscan be calculatedwith considerableaccuracyat specificationorany other set of flight conditions. The Iongitudinal-sy_netricaircraftcharacteristicshave been more accuratelydescribedthen the lateral-directional-antisymmetricset.
The aircraftride quality characteristicshave been calculatedusing thedescribeddata set and have been presented in referenceI, pages 56 and 57.
52
Flight control simulatorFlight-test-derived frequency response .........
Analytical model actuator, K + 50
3.0- K = 1.93 deg/volt
-154.4 (s-205.5)Fitted transfer function (S + 82.8 ± j95.6)
0 I I I I I I I I ! I
0 2 4 6 8 lO
Frequency (f) - Hz
0 2 4 6 8 lO
0 _ J w I I I i i J ,'
Phase -20, _'__'w"A A
: angle ____A
deg -40
-60
Figure 25. - Comparisonof analyticalSMCS actuatormodels to flightand simulatortest data.
53
150
Phase 1O0 langle k
50
deg :
0_" 2 4 6 8 lO
-SO Frequency (f) - Hz I "'
12
-- Flight test data
, ( 50------Analytical data _-_--_) actuator
I n z (S-205.5)_cv 8 ......Analytical data, -154.4 (S +82.8 .+-J95.6)actuator,
Test derived
grad
I _*,
4 X't_..
!
00 2 4 6 8 I0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)Wt = 119 296kg (263 000 lbs),A= 65 °
Figure 26. - Effect of test-derivedSMCS actuatormodel, frequencyresponseof normal load factor at FS 571.5 (225)due to SMCS
vane deflection,SCAS on, SMCS on.
54
400
F\300 /
I l
Phase 200 I x
"" angle I \/_._z,_.I i \
100 u I \. deg / _ \" 0 I . I * I I
2 4 6_ 8 JO
-100 - Frequency (f) - Hz
-200 -_" ....
Flight test data
Analytical data10
m
In]g I
rad ,_1 _ I4 I \ ;I _JIII
2 I l, \/I _
0 _* a_ , a , * * a ,0 2. 4 6 8 0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = I19 296kg (263 000 lbs),A=65 °
Figure 27. - Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)due to SMCS vane
deflection,SCAS off, SMCS off.
55
Figure 28. - Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225)due to SMCS vane
deflection,SCAS on, SMCS off.56
200 '
loo _ r\-. Phase 0 /"'_-_- _ _.,/ , _'._
angle 2 4 6_ L ' /-lO
de_ -100 _re_uen_If) Hz "7" _ J-200 "" ""
-300
10
Flight test data)
_SMCS 1at gain 1.5Analytical data
8
_Y I 6 [_cv I I
Ig
rad 4 !tI
t I |I i \2 I iI I \
00 2 4 6 8 0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = 119 296kg (263 000 ]b),A= 65°
Figure 29. Comparisonof flight test and analyticaldata, frequencyresponseof lateralload factor at FS 571.5 (225) due to SMCS vane
deflection,SCAS on, SMCS on.
57
FORWARD SMCS SENSOR PACKAGE RELOCATION
On the whole, the SMCS has worked well in improvingthe ride qualityofthe low-altitude,high-speedflightregime operatingat cockpitsettinggainsof i.5 in the verticaland 1.5 in the lateral. The lateralaxis performance,however,has been below that of the vertical.
As part of the ongoing investigationto determinehow to improvethelateralSMCS performance,the data shown in figure 30 were obtained. Thedata shownare power spectraldensity (PSD)plots of the pilot stationlateralaccelerationresponsedue to turbulencemeasuredduring flight 1-20 whileflying at M = 0.80 at about 305 meters (i000 feet) altitude. The data showthat the SMCS, with the lateralgains set at i.5, significantlyreducesthekey peak response at 4.5 Hz, slightlymodifiesa secondpeak at 6 Hz, andincreasesthe responsesignificantlyat 7 Hz. The net effect is an improvementin ride qualitybut not a large one. When the pilot increasedthe lateralgainto 2.2, therewas a dramatic increaseof the approximately7 Hz responsetoa level which was feltby the crew to be not acceptableeven thoughthe 4.5and 6 Hz responseswere furtherreduced.
The comparisonof the flight-testand analyticallateralaccelerationfrequencyresponsesdue to vane excitationshown in figure 29 d_nonstratedthat the analyticalmodel could reproducethe essenceof 7 Hz increasedresponse.A study of the analysisresults indicatedthat the two peak responsesat 4.5and 6 Hz are due to fuselagesidebending componentsin these modes while theresponseat 7 Hz is the fuselagesecond torsionalmode. Figure 31 will assistin showinghow the responsephenomenonobservedoccurs. Key in the analysisis the locationof the SMCS sensor packageat fuselagestationFS 571.5 (225),WL 142.24 (56)and BP 60.96 (24). Both the vertical and lateralaccelerometersare located in this package. As the vanes are differentiallydeflected,a sideforce and a torqueare created. The lateralaccelerometersees lateralaccelerationdue to both the side force and torque. When the lateralaccelera-tion signalsare sent throughthe SMCS, the side bending-ir_lucedsignalsareproperly phasedbut the torsion-inducedsignalsare adverselyphased,resultingin a reducedgain margin of the 7 Hz mode. It is also importantto note thatthe verticalaccelerometeralso sees the torque-inducedmotion,and undesirableSMCS symmetricalvane motionsare causedby lateralSMCS operation. Data,however,have shown this not to be a large influence.
Analyses shown in figure 32 indicatethat observedadversetorsioncoupling,as well as the coupling into the verticalaxis, could be eliminatedor attenuatedby relocatingthe SMCS sensorpackageclose to the fuselagecenterlineand near the elasticaxis. Since there would be no lateralmoment
arm, the couplingto the vertical axis would actuallyresult in the torsionalsignal phasingbeing favorable.
58
SMCS off
SMCS on, vert gain 1.5,1at gain 1.5
• o..... SMCS on, vert gain 1.5,1at gain 2.2
Flight 1-20
M = 0.80, air = 305 m (lO00 ft)
A= 65°
Figure 30. - Effect of SMCS lateralgain on power spectraldensityof lateralload factor at pilot stationFS 746.8(294).
59
SMCS nz
sensor package torsion I
An iIonat F$ 571.5(2251 ' Ytorsold location
' 60.96 c£ -_(24. in.),
i142.24 cm__Estimated
(56 in.) //'elastic
axis
_-_ 16.05 cm(6.32 in.)
I 9..62 in.)
ISMCS
ISMCS vane
:sensor package Resultant 1
at FS 515,6(203) Force vectors" " due to vanenew location
deflection
Looking toward rearfrom nose
Figure 31. - SMCS sensorpackagelocationsand couplingcharacteristics.
SMCSoff
8 _k _ SMCS on -" ny 1 old sensor
locationI lat gain 3.0
g 6rad SMCS on -
new sensorlocation
4 FI lat gain 3.0'I
"q
O
o
2
%eO eS
•If •
0O 2 4 6 8 lO
Frequency (f) - Hz
I. • I I I .... I I ,I
0 I0 20 30 40 50 G0
Frequency (w) - rad/sec
M-- 0.85, alt = 762 m (2500 ft)
Wt = ll9 387kg (263 000 Ib), A = 65°
Figure 32. - Effect of SMCS sensorpackage locationon lateralload factorat FS 571.5 (225)due to differentialvane deflection,analyticaldata.
61
The advantagesof moving the SMCS packagemay be sLm_narizedas follows-
(i) Higher lateralgains could be used to improvelateralride quality.
(2) Higher effectiveverticalgains (dueto a farther forwardsensorlocation)would improvevertical ride quality.
(3) The couplingof lateralvane inputsto the verticalaxis would beeliminated.
As substantiatedby both pilot co_ent and flight-testdata, the originalSMCS considerablyimprovedthe ride quality. However, the gains in the systemwere limitedto values below those originallyintendedby the couplingdescribedherein and not by systemmaximum capability. Based on the describedadvances,action was taken to move the forwardSMCS sensor packageto the new locationat FS 515.6 (203),WL 19.36 (7.62)and BP 16.05 (6.32). This relocationwasfirst accomplishedon A/C-I and then on A/C'2; this modificationwas neveraccomplishedon A/C-3. A/C-4 has the sensorpackage at the new location.
SMCS stabilitytests were initiallyperformedat high altitudeto evaluatethe forwardSMCS sensorpackage relocationeffects. The flight condition/configurationwas M = 0.85, altitude 6096meters (20 000 feet) and weight119 297 kilograms(263 000 pounds).
The First fuselageverticalbending-modedamping obtainedfrom the pitch-pulse transientdata is shown in figure 33 and is seen to be a linear functionof the S_4CSgain. All other modes were stable at the indicatedgain conditions.The first fuselagebending-modedampingwith the SMCS forwardaccelerometerrelocated (flight1-41) is comparedwith the resultswith the accelerometerin its previouslocation (flight1-7). The mode dampingappears to be signifi-
cantly larger at the higher gains with the new sensor location.
The lateralbendingmodes were not stimulatedsignificantlyby the rudderpulses, so that similardampingcharacteristicscould not be obtained. However,the lateralSMCS was stable for all values of gain tested (maxim_ cockpitknob settingof 6). Resultsobtainedwith the previous sensor location(flight1-7) showedthe SMCS to be unstableat a settingof 6.
• Followingthe high-altitudetest, stabilitytests were conductedat lowaltitude. The initialflight condition/configurationwas M = 0.85, altitude914.4 meters (3,000feet), A = 65 degrees,and weight 119 297 kilograms(263 000 pounds). The excitationswere horizontaltail and rudder pulses.First, each axis was tested to a maximum gain for that axis (V-gain= 3.0 andL-gain= 3.0, respectively). Followingthis, tests were conductedwith the
62
Legend: "--O-----O-- _
Flight 1-41 Flight I-7Mach = 0.85 Mach= 0.85
A = 65 degrees A = 65 degreesAir - 6096 m Alt = 6096 m
(20000ft) (200ooft)Wt = I19 297 kg Wt = 131 544 kg
(263oooIb) (290ooo Ib)Relocated SMCS Original SMCSsensor sensor
0.3
JI
J.E JO_
E f0.2--
o jI=
"O
_o.1- /7 ._ "
12 (
L !: 0 ' !, ,o 2 4 6
Vertical SMCS gain (cockpit knob set)
Figure 33. - First fuselageverticalbendingmode dampingversus SMCSgains for original and relocatedforwardsensorpackage.
63
SMCS operatingin each axis simultaneously. The maximum combinedsettingwas V-gain and L-gain = 3.0. The qualityof the response data was such thatthe d_qlpingfactorcould not be numericallydetermined;however,these qualita-tive data indicatedthat the SMCS was stable.
Followingthe horizontaltail and rudder pulse-stabilitychecks, SMCSresponseto SMCS vane excitationevaluationswere made. Figure 34 containsthe vertical-axis-responsedata. Shown are-thenormal load factoratFS 515.6 (203) (sensorlocation)frequencyresponsedue to SMCS sy_netricvane deflectionsfor SMCS off and SMCS on at vertical gains of 1.5 and 1.8.
These data show good performancefor the verticalSMCS and are not signifi-cantly differentfrom the similardata of figure 24 for the SMCS forwardsensor packageat the original location.
Figure 35 shows the lateralload factor at FS 515.6 (203) (sensorlocation)frequencyresponsesdue to differentialSMCS vane deflectionsfor SMCS off andon at lateralgains of i.5 and 2.2. As shown, the large 5 Hz (approximate)first fuselagelateralbending mode was significantlyattenuatedat the highergain without the previouslynoted large adversetorsionalcouplingresponseat approximately7 Hz.
Figure 36 shows the lateralload factor at FS 746.8 (294) (pilotstation)frequencyresponsesdue to differentialSMCS vane deflectionsfor the SMCSoff and on at lateralgains of 1.5 and 2.2. As on the previousplot, the
large 5 Hz first fuselagelateralbendingmode was significantlyattenuatedat higher gains. The 7 Hz response,however, shows a slight increaseinmagnitudeover the zero gain response.
Additionaldata similarto that just described,but at a higherweight
condition,are shown in figures 37 and 38. These data indicateless attenua-tion of the 5 Hz mode and more excitationof the next higher frequencymode
peak.
Rememberingthat the lateralnominal gain setting is i.5, all of thesedata indicateda substantialnet improvementin lateralload-factorresponsedue to the SMCS.
To furthercheck out the relocatedforward sensorpackage,tests were
conductedat the off-designconditionat M = 0.55, altitude762 meters(2500 feet),A : 55 degrees. Figure 39 presentsthe normal load factoratFS 515.6 (203)frequencyresponsedue to sy_aetricSMCS vane deflectionforSMCS off and on at vertical gains 2.5 and 3.0. These data show good perform-ance for the vertical SMCS at this off-designcondition.
64
SCAS only
SCAS + SMCS, vert gain 1.5...............SCAS + SMCS, vert gain 1.8
.I I I I I , I I
: 0 I0 20 30 40 50 , 60
Frequency (w) - rad/sec
Fit 2-20, runs 9.1,9.2,9.3M = 0.85, alt = 914m (3000 ft)
Wt = 129 502kg (284 500 lbs), A= 65°
Figure 34. - SMCS vertical axis performancewith relocatedforwardsensorpackage, frequencyresponseof normal load factor at
FS 515.6 (203)due to S_CS vane deflection,case i.
65
SCAS only
SCAS + SMCS, lat gain 1.5............SCAS + SMCS, lat gain 2.2
16
12
8g
r-_
/ \4 / ._ .
/ m
00 2 4 6 8 10
Frequency (f) - Hz
I, I I I I I I
0 I0 20 30 40 50 60
Frequency (w) - rad/secf
M = 0.85, alt = 914m (3000 ft)
Wt = 122 018kg (269 000 Ibs), A= 65°
Figure 35. - SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralload factor
at FS 515.6 (203)due to SMCS differentialvane deflection,case I.
66
r
SCAS 0nlySCAS + SMCS, lat gain 1.5
............SCAS + SMCS, lat gain 2.2
-- I0
8
cv I
0 , . .--- '. • _"Ii ''" t. I I ! IO 2 4 6 8 I0
Frequency (f) - Hz
I I I I I ' I
0 I0 20 30 40 50 60
"_ Frequency iW) - rad/sec
M = 0.85, alt = 914m (3000 ft)
Wt = 122 018kg (269 000 Ib), A= 65°
l:igure36. S_S lateral axis performance with relocated forward sensorpackage, frequency response of lateral load factor at FS 746.8 (294)
due to SMCS differential vane deflection, case i.
67
SCAS o_ly' SCAS + SMCS, lat gain .75
SCAS + SMCS, lat gain 1.5+ + + + + +SCAS + SMCS, lat gain 3.0
16
12
CV I[
8g
rad |'I
.
4
• ..I.4
0 s i i I I I I I I
0 2 4 6 8 I0
Frequency (f) - Hz
I I I I I I I0 I0 20 30 40 50 60
Frequency (o_)- rad/sec
M = 0.85, alt = 487.m (1600 ft) '
Wt = 139 709,kg (308.000 Ib), A = 65°
Figure 37. - SMCS lateralaxis performancewith relocatedforwardsensorpackage, frequencyresponseof lateralload factor at FS 515.6 (203)
due to SMCS differentialvane deflection,case 2.
68
SCAS only
SCAS + SMCS, fat,gain .75
SCAS + SMCS, lat gain 1.5
+ + + + + SCAS + SMCS, fat gain 3.0
12
lO
8
ny .
.6cv . .
6 . .. .
g 44
rad .
4 _"4
4-2 /,
@
0 ' I I I I I I I I
0 2 4 6 8 lO
Frequency (f) - Hz
I I I I I , I i
"- 0 l0 20 30 40 50! 60
Frequency (0_)-rad/sec
M = 0.85, alt =487m (1600 ft)Wt = 139 709 kg (308000 Ib), A = 65°
Figure 38. SMCS lateral axis performance with relocated forward sensor
package, frequency response of lateral load factor at FS 746.8 _294)due to SMCS differential vane deflection, case 2.
69
SCAS onlySCAS + SMCS, vert gain 2.5
....... SCAS + SMCS, vert gain 3.0
12
8
4
g
rad
00 2 4 6 8 0
Frequency (f) - Hz
I I I I I I I0 I0 20 30 40 50 60
Frequency(_) - rad/sec
M : 0.55, alt : 762 m (2500 ft)
Wt = 132 451 kg (292 000 Ibs),A: 55°
Figure 39. - SMCS verticalaxis performancewith relocatedforwardsensorpackage, frequencyresponseof normal load factorat FS 515.6 (203)
due to SMCS vane deflection,case 2.
70
Figure40 shows the lateral load factor at F$ 515.6 (203) frequencyresponsedue to differentialdeflectionof the SMCS vanes for the SMCS offand on at lateralgains of 2.5 and 3.7. As shown, the large 5 Hz firstlateralside-bendingmode was significantlyattenuatedat the higher gainswithoutthe previouslynoted adversetorsionalresponse at approximately7 Hz.
"" In additionto the precedingfrequencyresponse data, SMCS performancedata in turbulencewere obtained to evaluatethe relocatedforwardSMCS
sensor package. Figure 41 is a power spectraldensity plot of the pilot station(FS 746.8 (294))vertical load-factorresponsedue to ttmbulencewith the SMCSoff and on at vertical gains of 1.5 and 1.9. As shown, the SMCS was veryeffectivein attenuatingthe first fuselageverticalbending response.
The power spectraldensityof the load-factorresponsewas normalized
to unit root mean square (RMS) gust intensityCOw_) as derivedfrom the angleof attack (_ vane) data. This is not an accurate_techniquebut is the best
availablein absenceof gust boom data. A similarnormalization(OVg) has beenattemptedfor the lateralaxis data using sideslip (B vane) data.
Figure 42 shows the lateralload-factorresponseat the pilot stationwith the lateralSMCS gain zero but the vertical gains at 0, l.S, and 1.9.Since there is r_ mechanismfor the vertical axis SMCS motion to couple intothe lateralaxis, these data indicatethe level of repeatabilityof thelateraldata.
Figure 43 containsvertical load-factorresponse data at the pilot stationwith the SMCS off and on at vertical gains of 1.5 and 1.9, togetherwithlateralgains of i.5 and 2.2, respectively. Comparisonsof these data withthe data of figure 41 show that littleor no coupling of the lateralaxisactivity is evident in the vertical axis response.
Figure 44 shows the S_S performancein the lateralaxis at FS 515.6(203)with the SMCS off and lateralgains of i.5 and 2.2. These data showthat the 5 Hz (approximately30 radiansper second)first fuselagelateral
: bendingmode is significantlyattenuatedat both gain settings. However, atgain settingI.5, the secondfuselagelateralbending-modepeak response near6 Hz r_aainsabout the same in magnitudebut shifts slightlyupward in fre-quency. At lateralgain of 2.2, a significantincrease in magnitudedevelopswith this frequencyshift.
Similar data to figure 44 for the pilot stationare shown in figure 45.These data show the S Hz mode reductionbut indicatean increasedcouplingwith gain increaseof the highermodes.
Attention is directedto the power spectraldensity scales for figures144and 45. The data of figure 45 are a factorof I00 smallerthan figure 44.
71
SCAS only
SCAS + SMCS, fat gain 2.5
........ SCAS + SMCS, lat gain 3.7
2
g
rad
00 2 4 6 8 10
Frequency (f) - Hz
I, I ! a I m./___ e I
0 l0 20 30 40 50 60Frequency (OJ)- rad/sec
M = 0.55, alt = 762 m (2500 ft)
Wt = 129 276 kg (285 000 Ibs),A= 55 °
Figure 40. - SMCS lateralaxis performancewith relocatedforwardsensor
package,frequencyresponseof lateralload factor at FS 515.6 (203)due to SMCS differentialvane deflection,case 3.
72
Run ll.l vert gain = 0.0, lat gain = O.0------ Run II.2 vert gain = 1.5, lat gain = 0.0
........ Run ll.3 vert gain = 1.9, lat gain = 0.0SCAS on
200
xlO_ 6 "200
xlO-5
160
160 _nz
120
12o g2g2 (ft/sec) 2
(m/sec)2 "rad/secrad/sec 80
8O
|40 40
0 00 10 20 30 40 50 60
Frequency (_) - rad/sec
I I "1 I I • I :_ | , I I I I
0 2 4 6 8 I0
Frequency (f) - Hz
M = 0.85, alto152 m(500 ft)AGL
Wts 128 369kg(283 000 Ib),A = 65°
Figure 41. - SMCS performancewith relocatedsensor,PSD of verticalload factor at FS 746.8(294)- pilot station,case I.
73
Run II.I vert gain = 0.0, lat gain = 0.0
.... Run 11.2 vert gain = 1.5, lat gain = 0.0
..........Run ll.3 vert gain = 1.9, lat gain = 0.0i SCAS on
0.28xlO-6 .-
o.28x]O-5
0.24
0.24
0.20
O.20 _ny
o.160.16 2
2 gg (f t-TTsec) 2
(m/--_ec)2 "rad/sec
rad/sec O. 12
0.12 !o. 08 o. 08 I .
0.04 0.04 \_ _1_/ _\
V 'IQQ
e -foe
0 0 I I I I l .-0 ]0 20 30 40 50 60
Frequency (u_)- rad/secI I _L I I I I I • I IO 2 4 6 8 lO
Frequency (f) - Hz
M = 0.85, altu]52 m(500 ft)AGL
Wtm 128 369kg(283 000 Ib),A= 65 °
Figure 42. SMCS performance with relocated sensor, PSD of lateral
load factor at FS 746.8(294) pilot station,case i.
74
-- Run 11.1, vert gain = 0.0, 1at gain = 0.0.... Run 11.6, vert gain = 1.5, lat gain = 1.5............ Run 11.7, vert gain = 1.9, lat gain = 2.2
SCAS on
200
xl2oo
xlO-5
160 ....
160 _nz
120
120 g2
g2 (ft/sec) 2(m/sec)2 .rad/sec
rad/sec 808o
'140 i
40
0 00 Io 20 30 40 50 60
= Frequency (_) - rad/sec' I I I I I I I I'" I I I
0 2 4 6 8 10
Frequency (f) - Hz
M = 0.85, altB152m(500 ft)AGLWts 128 369kg(283 000 Ib),A= 65°
Figure 43. SMCS performancewith relocatedsensor PSD of verticalload factorat FS 746.8(294)- pilot station,case 2.
75
Run If.l, vert gain = 0.0, lat gain = 0.0
Run ll.4, vert gain = 0.0, lat gain = 1.5
.......... Run ll.5, vert gain = 0.0, lat gain = 2.2
SCAS on --28 ....
28 x
x10-5
244%
24" "eQeeee
..
20 "'
20 _m. : '__LIZ
O_v :;!2 16
2 16 g(ft/sec) _
(m/see) 2 rad/sec !ill i T,,,i
rad/sec 1212
P,I
/\ I:._
': "'" "'",; "...-_ - .-
0 0 I',-. I i .....0 I0 20. 30 40 50 60
Frequency ((_) - rad/secI I I, i_ I ,,, I , _ m.... I0 2 ...4 _ 6_. 8 I0
Frequency (f) - Hz
M = 0.85, alt--152m(500 ft)AGL
Wt =128 369 kg(283 000 Ib),A= 65 °_
Figure 44. - SMCS performancewith relocatedsensor,PSD of lateralload factor at FS 515.6(203)- sensor location,case i.
76
Run 11.I vert gain = O.0, lat gain O.O
------Run 11.4 vert gain = 0.0, lat gain 1.5
.......... Run II.5 vert gain = 0.0, fat gain 2.2SCAS on
0.28
xlO -6O.28xlO-5
_0.24
0.24
eege
._ e • 60.20 : .':.
2 (_2V • "._--g 2 - "'".Q ..(m/sec)rad/sec 2 -!I :
0.16 _ 2 _I| "',ft/sec) ,_.i_;I "
rad/sec :| [ •
o.12 :J I :O. 12 ;I I •
l %..:\ e
O.08 \_- '
o_::o.04 .._\0.04
: Y
•e
O O I0 10 20 30 40 50 60
Frequency (_) -rad/sec -L_ I, I . , .A " • -_ .' i I !0 2 4 6 8 lO
Frequency (f) - hz
M = 0.85, alt B152 m(500 ft) AGL
Wt =128 369kg(283 000 Ib),!%-- 65°
Figure 45. - SMCS performance with relocated sensor, PSD of lateral
load factor at FS 746.8(294) pilot station,case 2.
77
This large responsedifferencebetweentwo fuselagestationswhich arerelativelyclose is not reasonable. Figure 45 data are believed to be thedata in error. Subsequentto flight 1-56, the accelerometerat the pilotstationwas found to be unable to hold a calibrationand was replaced.Becausethe pilot stationresponseis of prime importancein ride qualityevaluations,it was requestedthat the data of flight 1-55 be rerun. Thiscould not be accomplishedin the schedule;however,the data of this flightdo have limitedvalue and are presentedhere for completeness.
Figures46 and 47 are similarto the data of figures44 and 45, respec-tively,but with the vertical SMCS on togetherwith the lateral SMCS.
All of the data in figures41 through47 were obtainedat an aircraftweight of approximately128 369 kilograms(283 000 pounds). Data similartofigures44 and 45 are shown in figures48 and 49, respectively,for a weightof approximately120 204 kilograms(265 000 pounds). Comparisonsof thesedata indicatemore adversehigh-frequencymode couplingwith increasinglateralgains for the lighterweight configuration.
It is concludedthat the frequencyresponsedata show a significantimprovementin pilot-stationresponsedue to relocationof the lateralaccelerometer. The PSD data of the lateralresponseatthe pilot station,however, still show considerablehigh-frequencymode excitation. Taking allevidenceinto account, the lateralSMCS still appearsto providea netlateralresponseimprovementfor the nominallateralgain of I.5.
TRUNCATED ANALYTICAL MODELS
As previouslymentioned,the flexibleanalyticalmodel of the B-I usedin SMCS designanalyses has been describedusing nomalized vibrationmodes
of the structure. The model used I0 symmetricand 12antiso_ymetricmodes inmost analyses. In de_ng a_]xrti.cg_l_modglstO supportre?ringbase simulatorstudies,however,it was found that this number of modes caused computerequT_m-e_iYYequirementsto become excessive. The challengewas to retain theaccurateaeroelasticimpacton short-periodand Dutch-rollcharacteristicsand the main essenceof the structuraldynamicmotion as seen at the pilotstationand SCAS and SMCS sensors.
78
Run II.1, vert gain = O.0, fat gain = 0.OIII Run II.6, vert gain = 1.5, lat gain = 1.5
........Run ll.7, vert gain = 1.9, fat gain = 2.2-- SCAS on
_6 '"
._ 28- x I0-5
x 1024
24
2O
20 _ny
g2 16
(m/sec)'2 ft/sec)2
8
,\ r.4 4 .i
• ° I e
0 0 l0 20 30 40 50 60
' .,. Frequency (0Ji - rad/sec
I I I I " I I , I I I ' -'0 2 4 6 8 |0
J
Frequency (f) - Hz
M = 0.85, alt _152 m (500 ft) AGL
Wt _ 128 369 kg (283 000 Ib), A = 65°.
Figure 46. - SMCS performance with relocated sensor, PSD of lateral
load factor at FS 515.6 (203) - sensor location, case 2.
79
Run II.I, vert gain = O.O, lat gain = O.O
..m_ Run 11.6, vert gain = 1.5, lat gain = 1.5_..........Run ll.7, vert gain = 1.9, fat gain = 2.2::
• . SCAS on ;' .:0.28 , , ::
0.28 x 10 6 ." : :: .-5 :i!! "
xlO : :,.I i •
0.24 : ::,0.24 '.
0.20 jI :0 20
_ j •o.16 I I:2
2°'16 g _. I:g 22 (ft/sec) _,: li
(mlsec) radlsec I: ! i
rad/sec O. 12 I; _.0.12\-\'.
0.08 0.08 1_ij_il:I: I-
0.0/4 0.04 I.
I! ,,I:
%,G" Z ""• e
o o0 i-O 2-_ 30 ;40 ,50 60
.Frequenc, ) (oJ) - rad2._ecI I I I ,I I :I" _" I I ' - •
0 2 4 6 8 lO
Frequency (f) - Hz
M = 0.85, alt _ 152 m (500 ft) AGL
Wt = 128 369 kg (283 000 Ib), A = 65°
Figure 47. SMCS performance with relocated sensor, PSD of lateralload factor at FS 746.8 (294) - pilot station, case 3.
8O
Run 11.8, vert gain = 0.0, 1at gain = 0.0Run 11.11, vert gain = 0.0, 1at gain = 1.5
......... Run 11.12, vert gain = 0.0, 1at gain = 2.2SCAS on
28 x .........-6
"- 28 x lO ,.
-5 ":I0 "
24 !_
24 _ '.
y • I i!
2 16
g (ft/sec)(m/sec)2 rad/sec
rad/sec 12
12t
|.
8 \8 :\
00 10 20 30 40 50 60
Frequency (m) - rad/sec
I I I I I I I I I I !0 2 4 6 8 lO
Frequency (f) - Hz
M = 0.85, alt _ 152 m (500 ft) AGLWt _ 120 204 kg (265 000 Ib), A : 65°
Figure 48. - SMCS performance with relocated sensor, PSD of lateral
load factor at FS 515.6 (203) - sensor location, case 3.
81
Run 11.8, vert gain = O.0, lat gain = O.00.72 Run 11.11, vert gain = 0.O, lat gain = 1.5
x 10-5 0.64 _. .......... Run 11.12, vert gain = O.O, 1at gain = 2.2-
x 10-6 SCAS on ..oe
0.64 ""eo
e •
56 "'0 . m • o
0.56 - ••
O.48-:'.: •
0.48 - • :• •: •
0.40 - ,. :
0.40 - _ny
2 0.32-g
(m/sec) 2 IIo
rad/sec 2 0.24 - 1: - !_0.24- g 1: ::
(ft/sec) 2 I!rad/seCo.16 'li :_ :o.]6- I: I
j;
I: k0.08 - 0.08 \ :
0 _ 0 _ I I I0 10 20 30 40 50 60
Frequency (_) - rad/secI I I i I i I l I I Io .2 4 6 8 lO
Frequency (f) - Hz
M = 0.85, alt _ i52 m (5OO ft) AGL
jWt = 120 204 kg (265 000 lb),./L= 65°
Figure 49. - S_S performancewith relocatedsensor,PSD of lateralload factor at FS 746.8 (294) - pilot station, case 4.
82
l)ynamic analyses Were conducted to identify the key modes contribut.[ngto dynamic motion at the pilot station (the SMCSsensor was nearby) andSCAS sensors located at the nominal center-of-gravity (CG) location. Thetechnique will be illustrated using the longitudinal-symnetric case. rnthis instance, the normal acceleration frequency responses at the pilQtstation and SCASsensors, due to excitation by the horizontal tail, were
- employed as criteria to judge the degree of accuracy achieved with truncatedmodels.
Thus, given a flexibleaircraftdescribedby rigid-bodymodes plus anumber of structuraldynamicmodes, the structuraldynamicmode set istruncatedwhile none of the quasistaticaeroelasticeffectsof the eliminatedmodes are lost.
The data used to modify the aerodynamic derivatives are generated by themethod described herein and are identified as F_/Roratios, or [ IF/[ IN"The approach to generating the F/R ratios is as follows:
(1) Select sufficient modes to represent accurately the dynamiccharacteristics of the real system.
(2) Assume those modes not selectedfor the dynamicsimulationto bequasistatic.
(3) Excite the quasistaticmodes with the aerodynamicloadingsassociatedwith the dynamicmode,S control-surfacedisplacements,and rigid-bodydisplace-ments. The loads picked up in each mode are"determinedby solvingthesimultaneousmodal equations.
(4) The solutionsto the precedingequationsprovidethe informationnecessaryto calculatethe F/R ratios used to correctthe aerodynamicderiva-tives of the rigid-bodyand structuraldynamicmodes selectedfor thesimulation.
The equationsto be used in the exampleare as follows:
Rigid-BodyPlungeand Pitch ModesT
83
Symmetric Structural Mbdes, i = 1 to n
Before approaching the details of defining the specifics of the truncatedsimulation, a brief description: of the basic P/Rtechnique is discussed. Ifall of the n structural equations were eliminated but the aeroelastic impact onthe short period retained, the equations would appear as follows:
Rigid-Body Plunge and Pitch Mbdes
84
The F/R ratios st_wn, where all 10 structure modes are ilwolved, areobta ined as follows. The set of structural mode equations are set up asshown below and _1 through _10 are solved for each of the indicated unit
loadings of _, (_w/2Vo) &, (Vw/2Vo)q, and _, respectively.
] il i I i! I II I I. I , , ' i
As an exampleof how the F/R ratios are developed,considerthe derivative
CN_. From the normal force equationwhere _ = 1.0 and all other rigid-bodyvariablesare zero, the followingrelationshipis obtained.
nl through nlO were obtained from the simultaneoussolutionof the previous
e_uationset-_ora loadingCni_ for a unit value of _.
The expressionis reformedto obtain the F/R ratio.
85
The F/R ratios
are obtained in a similarmanner.
The extensionto the pitching-momentcoefficientsfollowsthe same lineof logic.
Considernow the longitudinal-symmetricequationswhere three structuralmodes (I, 3, and 5) of the i0 are dynamic,but the quasistaticflexibilityeffectsof the eliminatedmodes are retained. The bar over the coefficients
indicatesa modificationdue to the quasistaticeffectsof structuralmodes
2, 4, and 6 throughI0. TypicallyCN = ([CN]/[CN])CN, where CN is for a rigidvehicle.
Rigid-BodyPlungeand Pitch Modes
Symmetricstructuralmodes, i = i, 3, 5
86
For mode 1 (typical of modes 3 and 5 also)
• The F/R ratio correctionsfor these equationsare obtained in the follow-
_ng manner. Ass_ning that structuralmodes I, 3, .and5 will be dynamic,thesimultaneousequationsto be solved for unit loadingsappear as follows:
87
The F/P, ratio data for CN and Cm coefficients,given the solutionof theseequationsfor unit loadingsof _, (_w/2Vo)&, (_w/2Vo)q, and 6, are
obtainedas previouslyexplained. The F/R ratio corrections,to CNni, CN_i,
Cmni, and Cm_. are obtained as follows:l
where 72, 74 and 76 through710 are obtained from the simultaneoussolutionof the previous equationset for unit 71 loading.
c_,zJ c_,<,and
c._3 c_ s
are obtained for unit loadingsof 73 and 75, respectively.
m m
Ccnq' i Cm_s , and -----.Cm_$
Cm_Zi CrowI Cm_/,_
,r"
obtainedusing the pitchingmoment___elationshipsand the 7i solutionsforare
the unit loadingof 71, 73, and _5' respectively.
The
data are obtained in a similarmanner for unit loadingsof (_i/Vo),(_3/Vo),and(_5/Vo).
88
'llm ratios for modifyingthe coefficientsof the dynamicstructuralmodes
are obtainedusing the n2, n4, and n6 throughql0 responsesof the simultaneous
equationsfor unit loadingsof _I' _3' 775'(_i/VO)'(_3/V°)'and (@5/Vo).Typ i ca1] y :
Cn,._.c_,n24+C_,n.7?.Cn,n,o_,O-- %,_,= I 4-__I
cn,n,
Where n2, n4, and n6 through nl0 are from unit nl loadings.
Where n2, n4, and n6 through nl0 are from unit (_l/VO) loadings.
It can be shown that all of the quasistatic aeroelastic information isin the truncated modal equations by using the following logic for a typical.rig id-body aerodynamic co efficient.
A B
f
(F/R)3-mode modified X (F/R)7_mode quasistatic = (F/R)10-mode quasi-system system static original
system
" The followingset of numbersfor a test case of the techniqueillustratesthe accuracyretained in the coefficients.
89
S
F/R F/RCoefficient from A from B
CN_ 0.752514 0.752519
CN& I.015199 1.015201
CNq 0.661091 0.661086
CN_ 0.620383 0.620382
CM_ 0.602100 0.602126
CM& 0.948420 0.948413
Q_ 0.732947 0.732943
CM6 0.588629 0.588754
Figure 50 illustratesthe degree of quasistaticand dynamicaeroelasticinformationretainedin the three structuralmodes plus correctionsset ofequations. Shown is the frequencyresponseplot of the normal accelerationatpilot stationand CG. Note that at zero frequency,the truncatedmodel responseoverlays exactlythe original 10-modemodel response, indicatingthat all ofthe quasistaticinformationof the 10-modemodel has been retained in thetruncatedmodel. Furthermore,the truncatedmodel dynamicresponse is a goodrepresentationof the 10-modemodel dynamicresponse.
Truncateddynamic equationsmay be developedfor the lateral-directional-antisymmetriccase in a manner similarto that shown here for the longitudinal-sy_netriccase.
ANALYSISOF ,SMCSVANEAERODYNAMIC INTERFERENCEEFFECT
During the developmentof the SMCS vane configuration,wind tunnel testswere conductedto determinethe aerodynamiccharacteristicsof the vane.Fairlyextensiveaerodynamicinterferenceeffectswere observed in the forceand moment data for both the longitudinaland lateralcases during componentbuilduptests. Reference1 containssome of these wind tunnel data andanalysisof their sources;refer to this referenceas backgroundfor thematerialto be presentedhere. It is the purposeof this sectionto reportthe resultsof an analyticalstudy made to assess the importanceof the vaneaerodynamicinterferenceeffectson the dynamicsof the aircraftresponse.
9O
0
0 10 20 30 40 50Frequency (_) - rad/sec
lO structural modes,
120 1 _H unsteady aero
1_-_I 1 3 structural modes + corrections,
pilot / _H quasistatic aero80
rad
4O
0
0 10 20 30 40 50
Frequency (_) - rad/sec
M = 0.85 S.L.
Medium weight, A= 65°
Figure 50. - Normal load factordue to horizontaltail frequencyresponsecon_arisonsof full and truncateddynamicanalyticalmodel, SCAS on.
91
The analysesconductedwere made in thefrequencydomainwherenormaland lateral load-factor responses at the vane station were calculated dueto SMCSvane oscillatory deflections at various frequencies. The main reasonfor doing this was that flight-test data existed for these frequency responsesagainst which to check the reality of the interference effects modeled.Another reason was that the basic analytical model was available to conductthis study; only the vane aerodynamic characteristics had to be reworked toinciude the aerodynamic interference effects estimated. The following para-graphs describe how this was done and discusses the resLtlts of the analysesmade. Finally, an evaluation is made of these interference factors relativeto the importance to future similar designs.
Discussions of reference 1 identified general areas where the forcescaused by vane deflections were acting. As shown in figure 51, they were on(1) vane itself, (2) forebody, (3) wing-aft body, and (4) empennage. As asimpi:ification, it was assumed that these forces acted at point locations ineach of the general areas identified. These point locations were determinedfrom the wind tunnel test data of forces and moments for various stages ofconfigurat ion buildup.
The SMCSvane alone force and moment coefficient data were determinedanalytically (reference 1) while the associated interference force and momentcoefficients (CN, Cm, Cy, Cg, Cn) were determined directly from the wind-tunneldata. The structural mode generalized forces were calculated for each modeknowing the forces acting at the points defined previously and the structuralmode deflections at these points (F¢i). However, before any of these datacould be added, the transport time lag effect from the vane to the point ofload impact had to be considered. In the frequency domain, this lag effectwas included for each interference load by mutfiplying by
-1vo
e
whereX is the distance(+aft) fromthevane stationto the pointin question,is the forcing frequency, and Vo the velocity.
TableVI showsthe rigid-bodyand structuralgeneralizedforcecoefficientfonmlationforthe SMCSvane,includingintereferenceeffectsforthelongitudinal-symmetriccase. The SMCSvanelateral-directional-antisymmetricdata,includinginterferenceeffects,were assembledin an analogousmanner.
The frequency response data of figures 22 through 24 and 27 through 29 are
repeatedhere in figures52through57. The normalloadfactorresponsedataof figures52 through54 showthatthe interferenceeffectsmodeledproducea
92
-- Horizontal tail
Vane
Wing-aftbody Forbody+ a L , + a R
(a) Normal forces
Vertical tail
Wing-aftbody
J
Forebody
(b) Lateral forces vane + _R' - _L
Figure 51. - Typicalvane-inducedinterferenceforces.
93
TABLE r1. - S\ICS \":\.\E EFFECTI\E\ESS 1\CLUD1\G 1\TERfERE\CE EFFECTS,LO\G1TUD1\.-li.-SY~I.'IETR1C C\SE, FRECUE\CY DO~l-U\
150 f_,_\ -f %
,oo:.-7,"".,\degPhase i/ / "..!_ /'_/// ":_ _.
angle 5G
/ i I I I ° I I i•.._"-i • --
o L ? 2 : 4! 6 8 . "lo• /_, ' : ". I
-50 _'I_l" "..°,i_' Frequency (f) - Hz "I
Flight test data
Analytical data, no vane aero interference
.......... Analytical data, with vane aero interference12 P,
I " "nz : :"
. a% •o
8 " :
g ": : •
:'" -'::_\ _
/// -,_. /•"JI "- _- -" -" -"",, ..'/" .. ._,:,€---/
• • °" /,' ° .- _.J:; - ,_ _.. _# -. ..0 ""_"_"I_'_i_'"I I I "" I I I I I '
0 2, 4 6 8 10
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = I19 296 kg (263 000 Ib), A= 65 °
Figure 52. - Effect of S_S vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)due to
SMCS vane deflection,SCAS off, SMCS off.
95
15°
Flight test data
Analytical data, no vane aero interference
..............Analytical data, with vane aero interference
12 . .."
"i :'
nz .. ::e • o •
_CV ' " ' '
• , e
8 " "
rad
4 •
oe° • •
", ..'"0 I I"" I I I I I i0 2 4 6 8 lO
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = ll9 296 kg (263 000 Ib), A= 65°
Figure 53. Effect of SMCS vane aerodynamicinterferences,frequencyresponseof normal load factor at FS 571.5 (225)due to
SMCS vane deflection,SCAS on, SMCS off.
96
•.C,.",xPhase 100 n _ ,, .-'" I °" -.l.._.."| ."" _,q_, :".
eoo • • _ • •
-. angle • ",,'i ; -" °°'_X " t _. ,* "*
deg 50 : I _' _'.\. : J/ _ .'IV'"
•e 2 ; 4: 6 8 "',. 0
"'" Frequency (f) - Hz "'_-50
Flight test dataAnalytical data, no vane aero interference
..............Analytical data, with vane aero interference :".:..
e el •
• ".12 ' '
: .# e
nz , ,
8 .... : ": "_
grad
o°°
4
00 2 4 6 8 O
Frequency (f) - Hz
M = 0.85, Air = 762 m (2500 ft)
Wt = I19 296 kg (263 000 Ib), A_= 65°
Figure 54. Effect of SMCS vane aerodynamic interferences, frequencyresponse of normal load factor at FS 571.5 (225) due to
SMCS vane deflection, SCAS on, SMCS on.
97
Flight test data
-- Analytical data, no vane aero interference................Analytical data, with vane aero interference
4OO
300 !"_ "/', :\ .
200 l _"" - _',r..--. ,,,,,_.
anglePhaseI00 / "J/"_ I_.
_; Frequency (f) - Hzdeg 0 "- ' ! I I • _la,... I I I 1
....... _vL-W_ 02 ; 4 6
-200-l O0 " %%_._ _ ..........j
owI i il I
lO
ji o
8
I6
rad
4k/
Ii:
I: _
I: t i
;.....;.-• "_, a-.-_" • n I I I00- "-- ..2 4 6 8 0
Frequency (f) - HzM = 0.85, alt = 762 m (2500 ft)Wt = 119 296 kg (263 000 lb), A= 65°
Figure 55. - Effect of SMCS vane aerodynamic interferences, frequencyresponse of lateral load factor at FS 571.5 (225) due to
SMCS vane deflection, SCAS off, SMCS off.
98
Flight test data
Analytical data, no vane interference
...............Analytical data, with vane aero interference
8
6grad
2
• ooS ••_' i_._ _
00 2 4 6 8 0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = I19 296 kg (263 000 Ib), A = 65°
Figure 56. - Effect of SMCS vane aerodynamic interferences, frequency
response of lateral load factor at FS 571.5 (225) due toSMCS vane deflection, SCAS on, SMCS off.
99
Flight test data
.... Analytical data, no vane interference
...... Analytical data, with vane aero interference
200
""____-,_lOO "_--_-_ _
Phase _ /-I lb,_ _.
angle- 0 _'______._ ,/ . .__'_,., 0
deg .... 2""_" 4 6_ _ ' _!Frequency (f) - Hz_._
-100 .
-200 % "-
-300 ...." 'lO
8
o,
6
g
rad _I _.
4t
2 _ _
_o_
00 2 4 6 8 I0
Frequency (f) - Hz
M = 0.85, alt = 762 m (2500 ft)
Wt = If9 296 kg (263 000 Ib), _A= 65 °
Figure 57. Effect of SMCS vane aerodynamicinterferences,frequencyresponseof lateralload factor at FS 571.5 (225)due to
SMCS vane deflection,SCAS on, SMCS on.
I00
significantdegradationof the analyticalto flight-test-datamatches inamplitudeand phase. A similar,but less severe degradationof the lateralload factoranalyticalto flight-testdata matches due to inclusionof theinterferencesis indicatedin figures 55 through 57.
Why better agreementswere not obtainedof analyticalresults (with inter-"" ferenceeffects included)to flight-testdata is difficultto assess. First,
however, it is to be noted that the interferencedata were obtainedusingstaticmeasurementtechniques,whereasthe flight-testdata were dynamic. Thevane on the 0.036-scaleforce model from which most of the componentbuildup
data were obtainedproduced small force readingsof questionableaccuracyforthis type of study (momentdata were judged more satisfactory). The 0.1-scaleforebodymodel was judged to produceacceptableforce and moment data. This,then, impliesthe need for more expensiveand larger scale models tested,using dynamictechniquesin order to supportaccurateanalyticalvane aero-dynamicinterferencestudies. The analyticalto flight-test-datacomparisonswithout interferenceeffects displayedin the referencedfiguresare quite
good, particularlyfor the normal load factors,and suggestthat such anexpensiveapproach is not warrantedto obtain satisfactoryaerodynamicdesigndata. A key conclusionreachedas a result of this interferenceanalysis isthat a static 0.l-scaleforebodymodel is adequate for obtainingvane aero-
dynamics data that includethe significantlocal fuselageinterferenceeffects;this model producedthe basic data used in obtainingthe best analyticalto flight-testdata match displayedin this report.
IMPACTOF SMCS ON SELECTEDLOADS
BACKGROUNDOF USING SMCS IN THE B-I FATIGUEANALYSIS
Becausethe SMCS systemwas designedto be fail-safe,not fail-operational,
the originalB-I structuraldesignconceptwas that the aircraftwould havefull structuralintegritywith or without the SMCS. That is to say, thestructuralloads for both designand fatigueanalysesof the air£ramewere
: to be done with the SCAS operativeand the SMCS inoperative. The B-I SCASwas designed as triple-redundant,fail-operational,and (forthe gust loadsanalysis)fully effectiveat all times.
During the B-I design development,studiesof the expected airframeloadswere accomplishedusing the availableweight, stiffness,and controlsystemsdata as it evolved. Severalyears into the program, it was found that filterchanges in the symmetricSCAS were having a considerableimpacton the forebodyfatiguespectrum. Furtheranalyses indicatedthat operatingthe SMCS mini-mized the effectsof the SCAS changesand gave forebodyloads of a more consis-
tant magnitude. Also, the lower statisticalload levels computedwith the
I01
SM(_ ;letirewere thoughtto be more representativeof those that would besee, i_ serviceusage. In view of these facts and with Air Force concurrence,it wCJsdecidedthat the B-I synmetricfatigueanalyses for the low-altitude
imnetrationmission segmentswould be accomplishedwith the SMCS operating.l:atigueanalysesfor all other flight conditionsand all design limit loadanalysescontinuedto be performedwith only the SCAS activated.
GUST LOADS ANALYSIS DESCRIFFION
As with the ride quality analysisdescribedearlier, the dynamicgustloads analysisperformedwas a generalizedmodal analysiswhere the equationsof motionwere solved in the frequencydomain. This analysisused 14 modes of
motion: plunge and pitch rigid-bodymodes and 12 symmetricfree-freenormalstructuralmodes. Also includedwere two active controlsurfaces: the hori-zontalstabilizerand the structuralmode controlvane.
For the B-I aircraftfatigueanalysis,the missions for the expectedserviceusage were each divided into segments. Flight conditionsrepresentingthe mission segmentswere selectedfor analysis. For each flight condition,staticaeroelastictrim loads (to be used as mean load levels)and gust
dynamicresponse incranentalloads were computed. Both the trim loads and gustincrementalloads were issued for fatigueand fracturemechanicsanalysesas distributedgrid loads. Along with the gust loads were the estimatedtimeto be spent at the flight conditionand a graph of the expectedload exceedancesper hour flight.
Aeroelasticloads for the steady-flighttrim conditionswere computed
using wind tunnel-derivednonlinearaerodynamicdata for the rigid andtheoreticallycomputedaeroelasticincrementloads. These static aeroelasticcalculationswere performedwith structuralinfluencecoefficientsfixed at apoint near the aircraftCG.
The gust response equationsof motion that were used are presentedasfollowsin matrix notation (ref. 6). See the appendixfor symbol definitions.
102
Frequencyresponsefunctionsof the generalizedcoordinates,q, wereobtainedby solvingthe equationsfor a unit sinusoidalgust inputat each ofi00 equallyspaced frequencies.
GUST L(I_DSSTRUCTURALMDDEL
A completeaircraft structuralmodel was derivedfrom structuralflexibil-ity influencecoefficientsgeneratedby the B-1 InternalLoads Group usingfinite-elementmethods. These SIC's,which were used for both staticand
dynamicanalyses,were for partiallybuckled skins so as to be representativeof the aircraftstiffnessat the design-limitload level. On the lifting
surfaces,the SIC points were arranged in streamwiserows to make it easy tocomputethe required slopesand deflectionsfor the aeroelasticanalyses.Figure 58 shows the arrangementof SIC poin_ts,and table VII presents thestructuraldegreesof freedomused in this model.
TABLE VII.- GUST LOADS MODEL STRUCTURALDEGREESOF FREEDOM
Motion Type
Component X Y Z @x @y @z
Fuselage,fixed wing and fairings 57Nacelle, includingengines 4 6 17 2 2 2Moveablewing 4 4 58llorizontalstabilizer 2 45
Total = 203 Z i0 I0 177 2 2 2
Normal elasticmodes, free from rigid-bodyplunge and pitch constraints,were computedby the method of reference 7, as shown in the followingequation.
where:-i T
103
'. GUST LOADS AERODYNAMICS
Aerodynamicgeneralizedforces were developedusing the DoubletLatticemethod with the wing-fuselage-horizontal-tailcombinationrun at one time.l:igure58 presentsa diagram of the aerodynamicpanel arrangement. Geometryadjustmentswere made to the theoreticalmodel to improvethe correlationof the zero-frequencystabilityderivativesand pressuredistributionswithwind-tunneltest results. One particularadjustmentwas to leave a small gapat the root of the horizontalstabilizer. At eachMach number,the width of
the gap was adjustedto get the best match of stabilizerCL_ and spanwiseloading. For the SMCS vane, the aerodynamicswere run separatelyusing thevane modeledwith 30 boxes (5 chordwiseand 6 spanwise)in the form of asymmetricinvertedV. These data were correlatedagainst the averagewind-tunnel test data. Generalizedvane forcesdue to modal motion and gust werethen added to those from the wing-fuselage-horizontalcombination. No attemptwas made to simulateany vortex wake producedby the vane.
Deflectionsused to define the structuralshape for aerodynamiccalcula-tions were selectedfrom the SIC points availablein the gust analysismodel. •It was found to be necessaryto review the shape of each mode to insurethatthe deflectionschosen defineda smooth shape with no sudden changesor rever-ses in slope. Points eliminatedrepresentedconcentratedmass it_ns andfairing-nacelledeflectionsthat were not representativeof the true wing-bodystreamwiseshape.
Generalizedforcesdue to modal, vane, and horizontal-tailmotions and aunit sinusoidalvertical gust were computedat zero frequencyand eightfrequenciesbetween zero and approximatelyII Hz. In the processof computing•the aerodynamics,the downwash inductionmatrices generatedwere retainedonmagnetic files so that they could be used again with differentmode shapes,thus savingconsiderablecomputingcost.
At each of the i00 solutionfrequenciesrequired,generalizedaerodynamicforceswere obtainedby spline-curvefits of real and imaginaryparts of thegeneralizedforces computedat the nine frequenciesas previouslydescribed.To improvethe qualityof the spline interpolationsof the gust forces,the °gust referencepoint was transferredto a point just forwardof the aircraftaerodynamiccenter. This transfertends to flatten the curves;i.e., reducesthe rate of oscillationof the functionswith frequency. After fitting,thegust referencepoint was transferredback to the nose of the aircraft. Forconvenience,the gust referencepoint is normalllyplaced at the nose of theaircraftso that when the solutionfrequencyresponse functionsare used tocompute time historiesdue to a discretegust input,the penetrationof thegust starts at time equalszero.
1.04
Fuselage station centimeters
500 lO00 1500 2000 2500 3000 3500 4000 4500I I I I I I I I I I
Inches
200 400 600 800 1000 1200 1400 1600 1800
Component Panel No. of boxes _____'._'_..'_"_*I*/
Forebody 1 6 _I_
Body + wing stub 2 40
Outer wing 3 6
Outer wing 4 30 *SIC points used to define aero shape
Wingtip 5 12 +SIC points not used to define aero shapeHorizontal 6 6
Horizontal 7 36
136
Y motion SIC points --
FRL . t t , T. . +
oFigure 58. Loads analysis SIC point locationsand DoubletLattice geometry.
l:inaladjustments were made by direct ratioing to bring the generalized
forces in the rigid-body modes, due to rigid-body motions and gust, into
agreement with the available wind tunnel measured values of CL_ and CM_.I:actorsdeveloped in the region of zero frequency were applied to the fullrange of solution frequencies.
GUST LOADS CONTROLSYSTI_V_
Descriptionsof the active control systems(ref. 8) used in this gustanalys[sare presentedin figure 59. Also shown is the method used to formthe transformationmatrix, IT],for control-systemsfeedback,relatingcontrol-surfacedeflectionsto the generalizedcoordinates.
LOAD METHOD DISOJSSION
Gust responseloads were computedat selectedstructuralstations. (Seefigure 60.) The mode displacementmethod was used at each frequencyof solu-tion to computedistributedloads at the SIC points. Shears,moments,andtorqueswere then computedby finite sL_mnationsof the loads or the loads
tilnesthe appropriatemoment arms. Figure 61 presents the matrix equationsused to performthe load computations. The mode displacementmethod finds aset of net externalloads that are equivalentto the loads requiredto holdthe structureinthe deflectedshape attained (ref. 3). To obtain accurateresults,a linear superpositionof the elasticmodes that are used in thesolutionmust give a good descriptionof the structuraldeflections. This
requiresthat at least severalmodes of vibrationsthat are primaryto eachstructuralcomponentbe included. Modes selected shouldnot be restricted infrequencyto the maximumFourier frequencyof the solution,but shouldbechosen to obtain all significantcontributionsto the structuraldeflections.
Althoughthe final loads for structuralanalysiswere issued as grid ormass point loads, it was necessaryto compute shears and moments in order to
keep the ntmber of items used in the load calculationsto a manageablesize.On each structuralcomponent,a referenceaxis and load stationswere estab-lished for computingthe shears,moments, and torques. The LOAD GEOM matrixcontainedone row for each load item. Basicallythe LOAD GEOM row elements_Jre: for a shear, ones in the columnsrequiredto select loads outboardof theload stationaxis, and for a moment or torque,the arms from the load stationor referenceaxis to each required load point.
To improvethe accuracyof the computedshears,moments, and torques,each load acting on a SIC point was consideredto he a pressure evenlydistributedover a load box around the point. For any load box cut by a loadstationaxis, only the box area outboardof the axis was consideredin
106
Symmetric SCAS at SL, Khp -- .36
Compensation Notch Low pass Servo Actuator
Degrees
Symmetric SMCS for M = .85 at SL, K_o= .565
Notch Shaping Washout Actuators
Degrees
The matrix representing the control systems in the solution equationswas formed as:
Gains x polynomials Mode defl.ectionsand slopesevaluated at S = j0_ at sen_or locations
7-n_c_ ec_ _,l 41
-nzc_ 0 .n_.c cuz /
Where: CG = FS 2649 cm (1043 in), CV = FS 516 cm (203 in), _'=
o
t:igure 59. Control systems for gust loads..
FS 2367 FS 2896
oc (932 (l 40)
FS 922 FS 1377 FS 1872 FS 2520 FS 3366 FS 4018 (1582)
(363) (542) (737) (992) (1325) Horizontal tail pivot
i I' I I' Pivot
Dimensions: cm (in.)
+S Forebody Aftbody
ZFB
, Yt FB I I _ iI II I I
Figure 60. - Structuralstations for gust loads analysis.
I
c
Point load at each mass point, by mode displacement method
Calculation of loads; shears, moments & torques
NLI_I NLIj N_I
Acceleration at point i
NMjI
Pitch rate at point i
=
i
" Control surface deflections Where: N = Number of mass pointsNM = Number of modes, rigid
_u = T(_J) ta) NEM = Number of elastic modes_V(_){ NLI = Number of S, M, & T load
2.aI Z_KIM NM_I iterns
Figure 61. - Load calculationequations.
109
computing t:he load. The ._ment and torsion arms for this-load were taken totile centroid o1 the outboard area.
The partitioningof the load boxes was computermechanizedby subdividinga box cut by a load stationaxis into I00 small boxes (i0 equal division on aside). Area ratio and centroid coordinatesof each subboxwere computedusingas data the coordinatesof the load box corner points. The area ratio wasthe ratio of the subboxarea to the load box area. Coordinatesof each subbox
centro[dwere then checked, and all those outboardof the load stationwereretained. The retained subboxarea ratios and area ratio momentswere accunu-lated to obtain the total elementvalues for the LOAD GEOM matrix.
Other load items computedwere load factorsat selectedstations,pitchFate and acceleration,and control-surfacedeflections.
Frequencyresponse functionsresultingfrom these load item calculationswere computerplotted for visual review and also saved on magnetic files foruse in computingthe load exceedancecurves and in-phasecomponentloadcond i t ions.
GUST STATISTICALLOAD CALCULATIONS
For each load item, the turbulenceresponse power spectrtm,_o(_),responsequantitiesA and No, and a frequencyof exceedanceof load curve werecomputed. The methods used were as presentedin reference9 and as follows:
2
o
3600 VN -
0
Ii0
The Von Karman continuousturbulencespectrumwas used with the gustscale L equal to 152.4 meters (500feet). Also used were the sea,level vertical
gust parametersfor low-levelcontourflying- Pl = 1.0, bI = 0.823 m/sec(2.70 ft/sec),P2 = 1 × 10-5, and b2 = 3.246 m/sec (10.65Tt/sec). In the cal-
_.. culationsto computeA and No, the requiredintegrationswere performedfromzero throughthe highestfrequencyof the analysis,approximatelyii Hz.
• The load level that could be expectedto be exceededone time per airplanelife in each mission segmentwas determinedfor each load item. This was done
to define a set of loads that could be used to developdistributedloadingconditionsfor the fatigueanalysis. These load values were read from the
exceedancecurve of each item at one over the total number of flighthours forthe mission segment. By nature,all of these loads were positivenumbers,and
althoughthey were consideredto representa load-cyclecondition,the phasingrelationshipsbetween the loads were unknown.
LOAD PHASING
The problemof load phasingwas solvedby applicationof the correlationcoefficientmethod to producewhat are referredto as in-phasecomponentloadconditions. Referencei0 developedand used the correlationcoefficientto
expressthe statisticalcorrelationbetweentwo gust responseparameters.
o
For this gust load analysis,correlationwas developedbetweenall ofthe load items. Correlationcoefficientswere not actuallycomputed;but a
matrix [B],the elementsof which were proportionalto Pij_i_j,was developed.- Here i and j indicateload item numbersand, thus, the corresponding[B] matrix
row and column numbers.
[B]= Real [H(_)] I_w(_)l[H_(_)I_]
iii
l'_chcolumn of [B]was then normalizedon the diagonal elementand
weightedw_th the correspondingload item expectedvalue.
where:
i=j
Columnsof the resulting [S]matrix representload conditionswhere thediagonal elementsare the load item expectedvalues and the off-diagonalelementsare the statisticallyin-phasecomponentsof the other load items.
CONDITIONMATCHING
To developdistributedgrid loads that matched the in-phasecomponentload conditionsthe mode displacementmethod was again used. As shown infigure 61, this method can be used to computedistributedloads for a givenset of generalizedcoordinates. For this matching problem,the shears,moments, and torquesat each load stationwere known, and the solutionhad tobe made for the generalizedcoordinates. The load generationmatrix wasdefined as:
Then the coordinatesand the shear,moment, and torque loads from the [S]matrix are related as:
LOAD] hSLOADS = GEN
112
_is equation is usually overdetermined,and the coordinatesare solvedfor in a least-squaressense. Due to the large differencesin magnitudeoftileshears and moments involved,the solutionresultsgenerallyprovidea poor
load match. To improvethe solutionquality,each load item was weightedso that it had equal magnitudeand, thus, equal significancein the solution.
. By multiplyingthroughby a diagonalmatrix of one over the loads, the followingequationwas obtained.
['°Ih1 = LOADS GEN
Taking one columnfrom the IS]matrix at a time, a generalleast-squaressolutionprogramwas used to solve for the generalizedcoordinates. Distributedi_int loads for the in-phasematchingconditionswere then formedby the modedtsplacementmethod.
EXAMPLE CONDITION LOAD RESULTS
Load resultsfor a B-I aircraftlow-altitudepenetrationconditionof
M '--0.85, altitude = SL, using sweep --67.5 degrees,and weight = 140 614 kilo-
grams (310 000 pounds)are presentedin table VIII. Also, presented infigures62 throughi00 are plots of the load-itemfrequency-responsefunctions,power spectrums,and load exceedancecurves for the items listed:
(i) Wing, WS 985 (387.6), bending moment
(2) Forebody,FS 1377 (542),bending moment
(3) Forebody, FS 2367 (932), bending moment
- (4) Normal load factor, nz, at CG
- (5) Normal load factor, nz, at pilot
(6) Delta (deflection) of horizontal stabilizer
(7) Delta (deflection)of SMCS vane (SMCSon)
113
'I'AI_I,I!VIII. I,OADCOMPARISON,SMCSOFF VERSUSSMCSON
lixpected loads, increments from trim, one o_ccurrence per 1000 hours• M = 0.85 at SL, wing sweep = 67.5°, wt = 140 616 kg (310 000 ib)
SCAS on
Loads
l,oad :items SMCSoff SMCSon
Shear WS 1341 (528) 38 330 (8 617) 39 149 (8 801)Bend. morn. 67 232 (49 588) 69 567 (51 310)Torsion ,' 3 243 (2 392) 3 246 (2 394)
Shear WS 985 (387.6) 80 406 (18 076) 78 622 (17 675)Bend. morn. I 278 977 (205 764) 280 389 (206 805)Torsion } 11 977 (8 834) 11 413 (8 418)Shear WS566 (223) 123 958 (27 867) 117 842 (26 492)Bend. morn. 698 319 (515 056) 681 015 (502 293)
Torsion _, 20 984 (15 477) 19.975 (14 733)Shear at wing pivot 171 710 (38 602) 155 567 (34 973)Roll morn. 649 463 (479 022 612 953 (452 093)
Pitch morn. 'i 1 355 627 (999 864) 1 280 991 (944 815)Shear at IrFroot 58 681 (13 192) 60 189 (13 531)
Roll morn. 237 236 (174 977) 238 159 (175 658)l'itchmorn. ,r 126 113 (93 017) 127 587 (94 104)
I:B morn. FS 922 (363) 381 315 (281 245) 188 692 (139 173)FB mom. FS 1377 (542) 1 060 369 (782 092) 349 942 (258 105)I;I_morn. FS 1872 (737) 1 848 893 (i 363 680) 364 677 (268 973)
I.B lllonl, tzS 2367 (932) 2 261 236 (1 667 810) 627 525 (462 841)AB morn. I-S2520 (992) 4 106 793 (3 029 030) 2 366 297 (I 745 300)
AB mom. FS 2896 (1140) 2 438 061 (i 798 230) 1 I00 992 (812 054)AB mom. FS 3366 (1325) 983 160 (725 145) 603 152 (444 864)
~
Nacelle Sz 85 953 (19 323) 91 366 (20 540) -
Anz at (:(; 0.956 0.949
An z at pilot 2.007 1.041 -horizontal, degrees 0.699 0.728
mode wine, degrees - 19.955
Stations, cm (in.)Shears, N (lb)bloments and torques, N-m (lb-ft)
1.14
-. N-m Ib-ft
m/sec ft/sec
14000. , ! ,
60 000 13ooo. I
,L,ooo. !! ' ''II I I
llO00. IIII ° I ,
'0000.40 000 M 9o00.8000.
'700o.6000, . ,
_ooo. ,/\_
20 000 4ooo.3ooo./f'_-j _' - \ I'2000. I :
1000. _ ._.
0 o0 20 40 ' 60
180.00
,oooo140.00 1120.00
!00.00
80.00 K
_.oo \ \P 40.0o\ _. \
H 20.00_ \ \
As o ',, \E -20.O0 \
"-40.00
-60.O0
-80.O0 X-loo.oo \
:: -120.00 _ !
-140.00
-160.00 t- " -180.00
20 q0 60
FREQUENCY, RAD IANSISECOND
Figure 62. SMCS off, wing bendingmoment frequencyresponse -WS 985 cm (387.6in.).
115
(:_(,b_ft)2m/sec/ ift/secrad/m rad/ft
I .OXlO+G9 , \
,o_- ,.ox,o._ \ Ill .................. _ III ....... -
\ T r'- 1 I I\ /I I zzz
r__z III ...........
P
V
,o,- , !lfl ",I HII
I]H.....'I.OXIO+05
I II I
0 .02 .0_ .06
SPAT|_ FREQUENCYRAD/FT
' ' ' ' ' 2'00 .04 .08 .12 .16 .
rad/m
Figure 63. SMCS off, wing bending moment response power spectrum -WS 985 cm (387.6 in.).
116
-. l.OXlO+03
l.OXlO+02
l.OXlO.01
EX I.OXIO +00CEEDANCES I.OXIO -01
/
H0UR
l.OXlO -02
l.OXlO-o3
I.OXIO-04
z 40000 80000 120000 160000 200000 240000 280000
13rr_NDING14C)IHEN'TAT NING STATION 387.6 lb-ft
• I ! i I
" 1O0000 .200000• 300000 .400000N-m
Figure 64. - SMCS off, wing bendingmoment exceedancesWS 985 cm (387.6inl).
117
N-m Ib-ft
m/sec ft/sec
150000
140000
6 x 105 13oooo120000
I10000
I00000
90000
4 x I05 i 8ooooTI VO000
oU 60000E
500001"
2 × lO_ 4oooo
ooo20000 ,-,
,oooo L/\0 o_ "----------__ _ _--------
0 20 40 60180.00
l_.oo14o.oo "\ 1120.00I00.00
_.0o
60.00 i
40.00 \
H 20.00 IA
_ o
-20. O0 t
-_o.oo I
-60.00-8o.ooi-I00.00
-120.00
-lqO.O0 \ I-)oo.oo r_ k_ I-180.00
20 40 60
FREQUENCY, RADIANSISECOND
Figure 65. SMCS off, forebodybendingmoment frequencyresponse -FS 1377 cm _542 in.).
II_
Figure 66. - SMCS off, forebodybendingmoment responsepower spectrum -FS 1377 cm (542 in.).
119
l.OXlO+02
l.OXlO.01
EX 1.0XI0 .00¢EEOAN
' CES 1.0Xl0 -01
/
H0UR
1.0Xl0 -02
1.0Xl0 -03
I.OXIO-04
4.0XIO+I 8.OXIO"
F'0REBDOYBO_[:)ING_ AT FUSELAGESTATIOXl_-t2 I b-ft
' 06 ' 06 ' 60.4 x 1 0.8 x 1 1.2 x lO N-m ""
Figure 67. SMCS off, forebodybendingmoment exceedances-FS 1377 cm (542in.).
120
N-m Ib-ft
-. m/see ft/sec_0000
300000
_0000
1.2 x lO6 eeoooo_0000
220000
M 2000O0
06 AG 180000
0.8 x 1 _ ,6oooo140000
120000I00000
20000
O- o_--_" __" _-___O 20 40 60
180.00
160.00 _
140.00
120.00
I00.00
80.00
6o.oo\
H 20.00AS 0E -20.00
-40.00 _ \_0.00 __
-80.00 _
-I00.00
-120.00
.. -140.00-160.00
-180.00 \20 40 60
FREQUENCY, RADIANS/SECOND
Figure 68. - SMCS off, forebodybendingmoment frequencyresponse -FS 2367 cm (932in.).
121
Figure 69. - SMCS of, forebody bending moment response power spectrum -FS 2367 cm (932 in.).
122
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X .OXlO+00
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< IllD IIIIIIIIII_IIIIIIIIIIIIIIIIIII IIIIIA IIIIIIIt111_1111111!1111111111 IIIII
Iliililiiliii i_'i illiliIii 'limeD' i'''i_ .o×,o-O,IIIlilflllllt_!ill II Illllll lit:::::::i:i:::lill_llllll iiiiiiiii'iliii_/ 'IIIIIIII,IIIIII[I_ ...... ,,,,,, ....
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!!.....OXIO-07
........................... iiiii]lill]l;lllll]llllllll]l]illlli_]l]'
,.ox,o-°,I!11111111111I I! IIIIIIIIIII ilqB.OXlO 1.6XlO' 2.4X1_ 06
F_Y B__J',(3INC_I'IOHENTAT FUS_AC_.,ESTATIONg_ 1b-ft
. ° # !
1.2 x 106 2.4 x 106 N-m
Figure 70. - SMCS off, forebodybendingmoment exceedances-FS 2367 cm (932in.).
123
g g
m/see ft/sec
•07000
.06_00
0.20 .06o0o.05500
•05000
M "04500
o. 15 _ .moooNI .0_::)00 |T
O. 10 u 03O00Q
E ._00
ooo/. \O.05 .ol_oo _.01000 W
.00500
0 .00 20 40 60
1130.00
160.00
140.00
120.00I00.00
80.00
60. O0
P 40.00 _
H 20.00AS 0E -20.00
-40.00
-60.O0 X,_-80. O0
-ioo.oo .!-120.00 \
, ooo-160.00
-180.00 1"_"0 20 40 60 -
FREQUENCY, RAD I ANS/SECOND
Figure 71. - SMCS off, normal load factor frequency response - aircraft CG.
[24
-3lO
! .OXlO-04
-4I0
0 .02 .0_ .06
SPATIAL FREQUENCY RAD/FT
f I I I I I
0 0.04 0.08 O. 12 O.16 0.20
rad/m
Figure 72. - SMCS off, normal load factor responsepower spectrum - aircraft C G.
125
l.OXlO +03iii
I
l.OXlO +02
\\
,ox,o.°, I\\
\
\E
X I.OXIO+00 \CE -\Eo \A \Nc \E \S 1.0XlO -01
/ \\
H0 \u \R
l.OXlO-02
\\
\\ ,
1.0XlO -03 \
I,OXlO -04 _0.2 0.4 0.6 0.8 1.0 1.2 1.4
LOADFACTORAT AIRCRAFT CO
Figure 73. SMCS off, normal load factor exceedances-aircraft CG.
1.20
g g
--. m/sec ft/sec
l.2 .35ooo
1.0 .3oooo
.25OOO0.8
MAG .20000N
0.6 IDU .15000E
0.4•I0000
O. 2 .05000 J L0 .0 0 20 40 60
180.00
160.00
140.00
120.00
I00.00
80.00
60.00
P 4o.oo!_ _.H 20.00
As o "X
E -20.00 _ _
_0.00
-60. O0
-80.O0
-mo.oo-120.00
-140.00 \-160.00
-180.00" " 0 20 40 60
FREQUENCY, RAD IANS/SECOND
Figure 74. - SMCS off, normal load factor frequencyresponse-pilot station.
127
rad_m rad/ft -_I.OXlO+00
i'ii ........!111I ........t1111I
,o° '_l"ll .....,.o,,o-o,i/lll
II i 11 I
i| Ill i[I Ill l
___ I Ill tII]]l
I ,! III I-1 I\ ./: tll ....
1o !1! /i t11
o, !1/![ Ill : ..... "_-
P
V _.... t-H_'_ ---2 " -
10 ,_ I11 .... T--,.............° !!1 /! _____-....-
=_ II I / ...........ill _ -
'3 III /
........ t-tl -- --------'----'--2 _---_._'-r_£t I I
t ] I __z: ::__rz_:TI=_-,= -o .02 .0,.t .o6
SPATIAL FREQUEN(;Y RAD/FT
I I I I I I
0 0.04 0.08 0.12 0.16 0.20
rad/m
Figure 75. SMCS off, normal load factorresponsepower spectrum -pilot station.
128
I.OXIO +03
l.OXlO+02
I.OXIO+01
EX l.OXlO +00CEEDANCES l.OXlO -Oj
/
H0UR
l.OXlO -02
l.OXlO -03
. _
I.OXlO-040.4 O.B i .2 1.6 2.0 2.4 2.8
LOAD FACTOR AT PILOT STATION
Figure 76. - SMCS off, normal load factor exceedancespilot station.
129
i
deg degm/sec ft/sec
.06500
O. 20 .06000.05500
.05000
O.15 .msoo#4 .04000A
O.10 I .030o0U
_ .0,2500
O.05 .0,.5o0.01000
•00500
0 .o _---..__ __0 20 40 60180.00
160.O0
140.00
!20.00 _ >
I00.o0
BO.O0
60.00
P 40.00 "_H 20.00AS 0E -20.O0
-40.00
-80.00
-80.00
,oooo / \ /-120.00-140.00
-I_::_).00 _,-IBO.O0
0 20 40 60
FREQUENCY. RAD IANS/SECOND
Figure 77. S_S off, delta horizontalstabilizerfrequencyresponse.
1.30
Figure 78. - SMCS off, delta horizontalstabilizerresponsepower spectrum.
131
I.OXIO+03
I.OXlO +02
I.OXIO+01
EX I.OXlO +00CEE
.DANCES l.OXlO -01
/
H0UR
l.OXlO -02
l.OXlO-03
F£gure 79. - SMCS off, delta horizontal stabilizer exceedances.
1.32
-. N-m Ib-ft
m/sec ft/sec
10000.
40 000 9ooo. I 1
8000.
30 000 M 7000. i
GA 6000.
lIN 5t700. _ /_
20 000 U 4000"://'_-'/3000. !_I0 000 2ooo. _..
1000.
0 o0 20 40 60
180.00
120.00
I00.00 " X
80.00
_°'°° \..... ",H _.oo \ \A ,,,
-2o. oo _-40.00
-60.00
-80.00 __-I0O.00
\ 1-140.00
-160.00
. . -180.00, 0 20 40
FREQUENCY, RADIANS/SECOND
Figure 80. - SMCS on, wing bendingmoment frequencyresponseWS 985 cm _387.6in.).
133
( N-m )2m/.secrad/m
(lb-ft )2ft/secrad/ft
J •OX 10+09
I.OXIO+08
oUTPUT
I OXIO+07P .5o
J .OXIO+06
\--I- -
\\
1/-f-
II -,
I' II-
--\ -
-1\
- -
-
-\
---.-
-\.,
- ~_.._- -- -
1\
-
\o .02 .04 .05
SPATIM. FRElU:t-.CY R/lD/FT
o 0.04 0.08 0.12
rad/m
0.16 0.20
134
rib~rc 81. - SMCS on, wing bending moment response power spectrum WS 985 cm (387.6 in.).
I. OXl0-°3
i
- - I.OXlO-°4100000 21_0000 300000
eENOmO.OMe_TATUmO STATION_37.6 Ib-ft
I I I I
lO0 000 200 000 30"0 000 400 000
N-m
Figure 82. - SMCS on,.wing bendingmoment exceedancesFS 985 cm _387.6in.).
135
N-m Ib-ft
m/sec ft/sec
40000.
35000.
1.5 x I0530000.
M _00.A
1.0 x 105 G
_ 20000.U
,,.o.Jr,t0.5 x 105 IOOOO. j_
_ooo. / "h
o0 0 20 40 60
180.00
i_o.oo _ ',.,
140.00 _,
120.00
lOO.O0
130.00 _,
\40.00
H 20.00A
°-20. O0
-40.00
-6°.00
-BO.O0
-I00.00-leo.oo \
J
-140.00 i _ k-160.00 _,-180.00
0 20 40 60
FREQUENCY, RAD I ANS/SECONO
Figure 83. - SMCSon, forebody bending moment frequency response -FS 1377 cm (542 in.).
136
(
2N-m
m/sec)rad/m
(lb-ft)2
ft/secrad/ft
I.OXIO+09
I.OXIO+08
oUTPUTp5o
1.0XIO+07
I.OXIO+06
l- - - .
1- -
IIIV\
- l- -I----1- - I-
,I-I-
I\y.-- - --
II II
1\1/
- -1-- - ,J..., -l--1 -
- -1\ -
- -
\
1\
\.-
-
1\1- -
- \1\
1\- - - ..
l-
I-- ~\I-- ''1,.0
o .02 .04
SPATlM... F"RECUN:V fW)/FT
.06
oI
0.04I
0.08. I
0.12
rad/m I
I
0.16 0.20
Figure 84. - SMCS on, forebody bending moment response power spectrum FS 1377 em (542 in.).
137
I. OXI0+03
~
,°
I. 0X10+02
1.0Xl0 +01
EX 1.0XI0 +00CEEDANcE \
• 5 1.0XlO -01
/
H0UR
1.0Xl0 -02
!
1.0Xl0 -03
i
I.ox10-°4 i100000 200000 ,.300000
FOF_EY3OYBO',OINGMOMENTAT _ STATION 542lb-ft
II I I I I
100 000 200 000 300 000 400 000 500 000
N-m
Figure 85. - SMCS on, forebodybendingmoment exceedances-FS 1377 cm (542 in.).
138
N-m Ib-ft
m/sec ft/sec"
2.0 x 105
I
45000.!
40000.
,.s× '°s . _oooo._°°"1 _NGk:_O0.
1.0 x 105 u_ 2oooo.D
0.5 x 105 loooo.
0 o _0 20 40 60
IBO.O0
160.00
140.00 L
120.00
I00.00
BO.O0
60.00
40.00
H 20.00AS 0
E -20.O0
_0.00
.oo 1
-_. O0
-I00.00
, oo ,\ /_.\-140.00
" -,_o.oo _., "\-,_.oo \
0 2O 40 60
-" FREQUENCY, RADIANS/SECOND
Figure 86. SMCS on, forebodybendingmoment frequencyresponse -FS 2367 cm (932in.).
139
,-
f-- - -f-
1\
J 1\ \\ II /~
-- f-
- - ---
1\ \.. - - . -
- - f-0- . -
"K- ._.-
\
-1- - - --
-f-
\-- -. f-
f- - - ~--
- \\-.
~r- -- .
- -
1.0XIO+06
I.OXIO+09
I.OXIO+oa
(Ib-ft)2
ft/secrad/ft
I.OXIO+ 10
ou~ I.OXIO+07
UT
P5o
610
( N-m )2m/secrad/m
0 .02 .O't .06
5PAT1~ FR£Cll.EN::Y RAD/FT -
'- , I I I I0 0.04 0.08 0.12 0.16 0.20
rad/m
Figure 87. - SMCS on, forebody bending moment response power spectrum FS 2367 cm (932 in.).
lila
I.OXIO+03
l.OXlO+02
1.0XI0+01
EX 1.0XI0+00CEEOANCES 1.0Xl0-01
I
N0UR
1.0Xl0-02
I.OXIO -03
I .OXlO-04100000 200000 300000 400000 ..500000 600000
Fg__.E_OYBO'-DING_ AT _ STATION 932 I b-ft
• i _- I ..... I !" 200 000 400 000 600 000 800 000 N-m
~n
Figure 88. - SMCS on, forebodybendingmoment exceedances-FS 2367 cm _932 in.).
141
, g g
m/sec ft/sec
.08000 _.
o.25 .o_o /!•07000
.06500
O. 20 .o6ooo. [Y5500
HA"05000
o.15 No ._oo [%/
I .04000Tu .03500
O. 10 _ .o_oo.02500
O.05 .o15oo \ _ X.oloooI "'xJ
.00500__ _ _-
0 .00 20 40 GO
180.00
160.00 1
140.00
120.00
I00.00
80.00
60.00
40.00
H 20.00A
-20. O0
-40.00
• "-60.00
-80.00 _N-I00.00
-120.00 -N._. ,,.
-1_o.oo _ %-160.00-18o.oo \ N
0 _ 40 GO
FREQUENCY. RADIANS/SECOND
Figure 89. - SMCS on, normal load factor frequencyresponse-aircraft CG.
1.42
.06.04.02o
1\-l- f-
- ~ - - - I- f-l- f-
1\
1\-- --~
2 1\-
- i- f-
-.-
1\- r-
II.i' I\U !\ IA
.-
-- ~
1\
II1\1\
i-
- 1\ i-iI,
~ \
\. - -- i-
- - -~ 1-. -1\ .. -- - - -
--1\ 1;< - -~. -~ ~
1\ i-
.. -t-.~ i-
" .... 1...-
1.0XIO-04
I.OXIO-O
oUT~ I.OXIO-03
T
P5o
(~2rad/ft
-210
-110
rad/m
SPATlf.L FREa..EN:Y RAD/FT
oI
0.04I
0.08 0.12 0.16 0.20
rad/m
Figure 90. - SMCS on, normal load factor response power spectrum aircraft CG.
i43
I.OXlO.03
I.OXIO+02
l.OXlO+01
EX l.OXlO+00CEEDANCES I.OXIO-01
I
H0UR
l.OXlO-02
l.OXlO-03
1.0XlO -°q
0.2 0.4 0.6 O.B 1.0 1.2 1.4- .
LOA_)FACT_ AT AIF_I:_-T CG
Figure 91. - SMCS on, normal load factor exceedances- aircraft CG.
144
g g
m/sec ft/sec
O. __00
.08000
.07000
O. 10 .o3ooo
ooo/ %.01000
0 .0 0 20 40 6OI_.00
160.00 I
I
140. O0 1 _ '120.00
I00.00J t i
80.00i i
60.OO
40.O0p t l , iH _ .00A t is o \/-,,,E __.00 ___.
-40.O0i
' \--60.00-80. O0
-100.00-120.00
-140. O0
-I_.00-180.00
" O 20 40 6O
FREOUENCY, RAD I ANS/SECONO
Figure 92. SMCS on, normal load factorfrequencyresponse - pilot station.
145
g 2 g 2m/sec ft/-sec
F-
10-1_ ' _
l.OXlO -02 /
-2 _ __i0
o
y i,OXlO_O3P ~_UT
P
O I
-3- /lo
i.OXlO -04
m
-410
o .02 .oq .06
SPATIAL FREQUENCY RADIFT
| I I I . .. I I
0 0.04 0.08 0.12 0.16 0.20
radim
Figure 93. - SMCS on, normal load factor responsepower spect_Jm - pilot station.
140
I.OXlO+03
\
I.OXIO+02
\
l.OXlO+01 \
\
i
\ JEX l.OXlO+00
[ \ 'Eo \A \Nc kE \S 1.0Xl0 -01
/ \\
H
o \uR
|.OXlO-02
l.OXlO-03
i
I .OXlO-°h0.2 0.4 0.6 0.8 1.0 1.2 1.4
LOAD FACTORAT PILOT STATION
Figure 94. - SMCS on, normal load factor exceedances-pilot station.
147
deg degm/sec ft/sec
0.1 0 - .o3ooo
o_ooA J i
// ".02600
O.08 .o24oo .•OP_2OO
.02000J
0.06 _ .o16oo,NO .olr_o : .
i .01400
0.04 u .oI_o ,E
.01000
.OOSO0 _ {
0.02 .oo6oo , I •.oo4oo I.0_0
0 .o "'_ _ , T_ ..-.L20 40 60180.00
160.00 _
140.00
120.00I00.00
130.00
60.00 1
o.oo !p -
20.O0 /H
A 0
[-_1.00
-40.00
-60.00-_.OO
-I00.00
_,_.oo \_-140.00-I_.00 \
-IBO.000 20 40
FREQUENCY, RADIANSISECOND
Figure 95. SMCS on, delta horizontalstabilizerfrequencyresponse.
1.48
/ deg _2 f deg ,_2m/sec _ft/sec/\ra--_m/ -r_--_T-
lo-1 _L_'X -
, x_
!
:-___\,OXlO_ I
o \ ---_v ......_--"........ _,.o×,o-O_ / Ip ..... / \
1o-5 s -- \ \ / \ I,.ox,o-°_ / - i "T
............. _.___-_-¥,=___--__-,-,_,_.-.-
,ox,oO, ! r.j.k-- ' -_--_-t:
10-7- _ .....I-
f.OXlO-°8
_=1==!:_--- -0 .o2 .1_._ .o6
,_PATI._L FREQUENCY RAD/F'T
L _l i I t
0 0.04 0.'08 0.12 0.16' 0.20rad/m
Figure 96. - SMCS on, delta horizontalstabilizerresponsepower spectrum.
149
150
EI.OXlO+00X
CEE0ANCE
J .OXIO-015
I
H0UR
I.OX1O-02 1\
oo...TA ~'ZCNT~ ST.6BILIZER. CE~S
Figure 97. - SMCS on, delta horizontal stabilizer exceedances.
m/sec ft/sec
1.7000
" - 1.6000
1.50001.4000 A
I.3000 /_
/I
1.20001'4 I. I000
1.ooo0
3 _ o.9ooo0.8000
g o.7ooo
0._00 I0.4000
I 0.3o000.2000
O. I000 _ __
0 o 40 60180.00
160.00 _,_
140.00 _
IL='O.O0 _ "-.,1oo.oo _ \
80.00
40.00PH c_.obA
o-;20. O0 N_-4O. 00
-60.00 "_-80.00 _
-I00.00
-120.00 _
-140.00-IBO.O0
-180.00'_ 0 2O 40 6O
FREQUENCY, RAOIANSISECOND
Figure 98. - SMCS on, delta mode controlvane frequencyresponse.
151
( deg _2 (deg _2\mlsec/ \ftlsec/rad/m rad/ft _.
I .OXlO.01 "
lO1
I.OXIO+Do
0I0
I .OXlO-010UTP
-] ulO T
PsO
I .OXlO-02
-210
l.OXlO -03
-3lO
0 .02 .0'-I• .06
SPATIAL FEQUEI',_Y' R,_/FT
I I I I I f
0 0.04 0.08 O. 12 O. 16 0.20
rad/m
Figure 99. - SMCS on, delta mode controlvaneresponsepower spectrum.
152
I.OXIO+03I_ Ill;....... [:::::11 I]I: !':_l[: ill]
[llj Ijjlljjj ijll ' _ill tljl
IIII ....................IIII]111 1tll I111 lilt1111 IIIIIIII IIII Itll I!ll
_llil lillilli Illi iili lili,.ox,o+o__t! t! tl IIli ltl i ii
_ ikii: _ iiiiiiii _; iiii iiii__!_iiii...................... IIIII I _1 11111111 I I I I I I I IItlM II111111 IIII ]111 Ill]til_ lltlllll_ ll,l Iljl t111illl' liltilil ltli Ilil IIII
,.=o+O,Illi' Ililli II Il! II::iiii::,.iiiiii!! .................llll TIlL z 1111
.................. III]IIII IIIIIIII Illl IIII
IIII \11111111 Ill! Ill! IIII
TITT i_lllill Illl litl IIIt
!!!! .... ,,!!]:lle IIII IIIi_III :-iiii .... :III ....I111 Ill I1"1 IIII IIII ]Illo IIII IIIll\lll I1tl Illl III1A IIII Illlt_ll lilt till IIitn Illi Illill_l fill IIII liltC
_,-°×,°-°',..! Illi1, I il! .it# ll:i iiiiiii[},i"iiii .... !!ii _:' iiii
IIII I1111111 IIII Ill[
IIII IIII111! IIII IIII illlH llll I lllll II It It Illl _. IIII0 IIII Iit11111 \ till Itll IIIIu IIII iliiilli ,ill! lill liliR
,.o×,o-°_Jill I1!1 N I J!! IJ......................... lil :
tr,,I III, _ i i ' ' i , ,\, I I II 1111 "ilil ........
'"' IIlitl ItlltlJt ll!i \ Itll IIIlli iillllti Itli _illl li
,.ox,o-°_II Ililti il "N. !!!till [Itzl[tl Tltl ._ tlll1111 IllllIlt IIll lIlt _ ill/
I111 IIIII!ll IIII II[I _ IIIIIIII IIIIIIII IIII t111 nb,.]llJill Illillll Iltl Ilil t f't.,.]
'. IIII IIIIIIII IIII IIII III1",.ox,o-_Ilil IIIIIIII III Jill III 7-_
4 8 12 18 20 _
" " DO-TA _ _ V,N_E, CE_:_E_S
Figure i00. SMCS on, delta mode controlvane exceedances.
153
The tabularload summaryresultsand the plots are presentedfor both_ICS off (gains= 0), and SMCS on. In eithercase, the SCAS was consideredto be operatingnormally.
As can be seen in the summarytable,the effectof the active SMCS was
to substantiallylower the fuselageforebodyexpectedbendingmoments. Alsoloweredwere the aft fuselagebendingmoments. The most obvious effectofthe &ICS, as can be seen in the forebody frequencyresponseand PSD plots, wasto considerablysuppressthe responseof the fuselage first bendingmode. Atthe same time, the operationof the SMCS caused increasedresponsesin someof the higher frequencystructuralmodes. The effectof this increasedmodal
activity,as seen in the fuselageand wing frequencyresponseplots, was tocause slight increasesin the wing tip, nacelle,and horizontaltail expectedloads.
Althoughthe effectof :theSMCS activitydramaticallyreducedthe magnitudeof the forebodyfatigueloads for low-altitudepenetration,the total forebodyfatiguespectrumswere not so drasticallyreduced. The SMCS was not usedduring the high-altitudecruise mission segments. The large numberof flighthours spent cruisingand the higher wing lift curve slope at a 25-degreewingsweep, cause the expectedloads to be relativelyhigh. Thus, with thereductionin the low-altitudepenetrationloads, the cruise conditionsbecamedominant in the forebodyfatigueload spectrum.
SMCSVANE EFFECT ON INLET/ENGINECHARACTERISTICS
]he objectiveof this section is to brieflydescribe and summarizeB-I
flight-testresultsthat identifyeffectsof ingestingvortices generatedby the SMCS vanes into the inlet. Although vorticesgeneratedby the SMCSvanes were ingestedfrequently,no engine incidentsrelatingto operationof the SMCS were identifiedduring the approximately1200 flighthours accu-mulated to date. This program includedmore than 200 flightswith threeaircraftand 29 engines. Wind-tunnelresultswith sub- and full-scalemodelswere summarizedin referenceI.
Portionsof specificflightswere dedicatedto demonstratingoperationalsuitabilityduring aircraftmaneuverswith the SMCS vanes deflected. Emphasiswas placed on exploringcombinationsof SMCS vanes deflectionangles and air-craft maneuversduringoperation at Mach 0.85. Effectson inlet total-pressurerecoveryand engine-facedistortionwere measuredby a 40-probe instrumentationmatrix at the inlet/engineaerodynamicinterfaceplane (AIP). Instrumentation,includingan automaticin-flightcalibrationsystem,and flight-testproceduresare described. Resultsare summarizedto doct:nentthis B-I experienceas anaid to futureprograms employingsimilarsystems. Nomenclatureassociatedwith this sectionmay be found in the appendix.
15,1
TEST DESCRIFFION
The B-I propulsionsystem is arrangedin two nacellesunder the fixedportionof the wing as shown in figure I01. Each nacellecontainstwo inde-pendenttwo-dimensional(2-D) externalcompressioninletsand two GeneralElectric F-101 afterburningturbofanengines. Relativelocationsbetween
- the SMCS vanes and inlet nacellesare also shown.
Most flight tests investigatinginlet characteristicswith the SMCS vanesdeflectedwere conductedwith the simulated,fixed-inletconfigurationshown
:infigure102. Ramp configurationfor both inletsconsistsof the initialtworamps set at 7 degrees. On the inboard inlet,the third ramp is set at5 degrees. In the outboardinlet, the third ramp is set at 9 degrees. Smalldifferencesbetween inboardand outboardramp configurationsreflect anattemptto maintain good performancecharacteristicsduringboth subsonicand supersonicoperation. The movable cowl lip is shown in its normal,takeoff,and landingpositions. Duct flaw area distributionsare shown infigure103. Maximumflow area is based on an averagethird ramp angle of7 degrees. Design flow area is shown for referenceand representssupersonicoperationwith a variablegeometry inlet.
Inlet boundary-layerair is removedthroughporous surfaceson the secondmovable ramp, throat panel, and small regionson the upper and lower end plates.The bleed air is collectedin two compartments. The air exits from the forwardcompartmentthroughfixed louversand from the aft compartmentthroughtwo-positiondoors. The doors are open above MachL_l.4and closedat lower speeds.Aft bleed doors were closed during all tests with the SMCS activated.
A bypass systan operatesat supersonicspeedsabove Mach 1.4 to match theinletsupply and enginedemand. The bypass doors open to compensateforreduced engineairflow such as occur on a hot day or during low-powersettings.The bypass doors remainedclosed during all tests with the SMCS activated.
The SMCS vane configurationis definedin referenceI. Flight-testpro-visions includeda black box located in the crew compartmentto drive the
2
SMCS vanes to a fixed deflectionangle. From this position,the vanes couldalso be driven at selectedfrequenciesto oscillate+i0 degrees. Using these
-- provisionsto generate full-scalevane deflections,flight tests were conductedto identifywake ingestioneffectsduring combinedaircraftmaneuversand enginethrottletransientsas measuredat the inlet/engineAIP.
155
Figure 101. - B-1 flight test aircraft, wings swept 65 degrees.
: Throat fixed link--
___'_- fimed
linkBypass
Ramp fixed link
Movable lip__." _:::::`_:_:_i_i_i_i_iiiii_ii!:iiii_!iiii!!ii_iii!iii_i;iiiiiiiiiii_iiiiiii_iiiiiiiiii!i_ii_i_iiiii!i_iiiii!ii_iiiii!i!iii)iiii_iii!i_ii%_iii;!iiiiii_iiiii_!i_
(takeoff/landing) __ _ . _____
"-- " _ _._ _ 2:2:i:i:fi22:2:i2)i:r:2:i:i:i:5:2!iii21211:i[:2:i:r:i:iiii:i21:2:2i:i:2:i::-_2:!:f:!:i:2:i:£_:2!i:ii1_:_:J:!:_:2:!:_ii!ii_:_:i:_:_:J:_:1!i_ii_i_:y:i:i:f:i:2:5:_::::::2:i:i:[:_:i:i:i!_:[£_i_2i_:_:[:J:_:i:_:J:[_i5!!_Ji_:_:i:_:2:J:J:!:2
Fixed throatY_ I°v:_::rI.o.,
Fixed b
_--BLC 1-[ two-position doorsAdjustable link(ground adjustable)
Figure 102. - B-I air inductionsystem.
157
18ooDuct flow area distribution
ll 000 ECI-8 Aerodynamic
interface _Plane_ l
Movable lip inlO000
I10258 cm 2
9000 (1590 in.2)((cm2 9258 cm2
(1435 in.2)8000i
4-1U
"0<[:,
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600C J500O
4ooc I i I a I _ I I I _ I I I l I l I i I80 lO0 12o 140 16o 180 200 220 240 260
in.
I I I I I |__ I I I I I I200 300 cm 400 500 600 700
Nacelle station (at duct _)
° Figure 103. - B-I inlet subsonic diffuserflow areadistribution.
FLIGHT-TEST INSTRUMENTATION
B-1 flLght-testaircraftwere instrumentedto measuremore than 1600
parametersof which approximately700 pertainedto the propulsionsystem.InstrL_entationat the inlet/engineAIP included40 dual-purposeprobes inboth engines in the left-handnacelleto measureboth the steady-stateanddynamic componentsof total pressure. Probeswere installedintegralwitheight engine-inletguide vanes, eachwith five probes locatedat the centerof equal areas. (seefigure 104.)
Computeddistortionparametersare sensitiveto errors in individual
total-pressuremeasur_aents. Several techniqueswere employedduring theflight-testprogramto minimizethese errors. High-response,differentialtransducersmanufacturedby Kulitewere referencedto a duct staticpressureupstreamof the inlet/engineAIP to maintain signalresolutionover a widerange of operatingconditions. The referencepressurewas measured by anaccurate,digital, absolutetransducerinstalledin a conditionedcompartmentin the aircraft. The referencesyst_iwas constructedwith sufficientvolume
and orificesto restrictrapid changesduring aircraftand/or inlet transients.Referencepressurewas sampledfour times per second.
An in-flightcalibrationsystemwas developedto update individualprobecalibrationsonce per minute throughouteach flight. A schematicof thissystem,using a three-waypneumaticvalve for each total-pressureprobe, isshown in figure 105. The valve alternatelysequencesthe back side of thetransducerfrom the normal referencepressure (operatemode) first, to acalibratepressureregulatedto a pressureapproximately5 pounds per squareinch above the referencepressure (calibratemode), and then to the samepressureseen by the front side of the transducer (zeromode) (hencethe nameZOC valves).
The calibrateand zero positionswere each held for 2 seconds,and theoperate positionwas held for the remaining56 seconds. In the zero position,the total pressurebeing sensed is routed to the back side of the transducer
: throughan infinitecoil, approximatedby a coiled line 20 feet in length.This provisionwas necessaryto preventreflectingwaves affectingthe frequencyresponseof the probes and was determinedempirically. Both coils and pneu-
-" matic valves, the latter groupedin gangs of five, are shown installedonthe engine in figure 104.
159
.....Q\o
lguide vanes
Total-pressure probe
Kul ite AlP probes359° aO
IView looking aft 359° aO
I
80°
3au
71°
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Engine
B-] j nlet/engine aerOd)l1amicEngine 2
tntcrfacc plane instrumentation.
Fuselage Engine compartment Inlet-- ----I i -II I l II I I l I
zocI' ' m '! Icalibrate Pneumatic ICal + 15 psi I ' ! I _ . 'I actuator
I I Ii l Zero "_ i[ I II
l uperate _ I
Cal I. II ZOC valve I
R2+ 5 psiI I I i • Ii I I
I I _.Infinite coil I Engine I
I R2 . I I I airflow I
Inletl'I I t__ "i I! I Ii
static I ! .AIP Drobej I , ,-
-- I _ ....... , ,_ ___j_-..... Si%n-al--co--ndT ti-one'-r............ I
I P°werII 7.5v supply IICrew station I _. [
i I Cal, zero _s I lTape
rec°rdeff ICBW Operate. _x2__._.j I lHP F _._j_ - III.__R2>10 ps!aTM _-- filter I
' ; R2_ O.4Hz I16 ips I L • Normalize
,.--I' ' iiI "_J ILP [- _rans_u_er 'A/D I LP fi1ter,
J 16 spsi_ filter,0.4Hz II 12HzI_ l' I II I L I
MV multiplexer I
_ Figure 105. - B-I flight-testAIP data acquisitionand signal conditioningsystem.
lhlet_naticvalves were actuatedby a pressure sourceregulatedto 15 poundsper square inchabove the calibratedpressure. Provisionswere includedtovent trapped pressures within the valve to either the engine compartment(ambient pressure) or an engine-face static pressure, depending on the operat-ing condition. Vent pressures were required to be the lowest pressure in thesystem at all times. Selection of vent pressure was accomplished by a crewcompartment switch. A pressurized nitrogen bottle was used as a pressuresource for the actuate and calibrate tanks.
Signal conditioningfor the transduceroutput is also shown in figure 105.l.hcbsignalis initiallypassed througha variable amplifierthat is used toestablisha full-scalerange for the steady-statecomponentat approximately24 pounds per square inch. Zero output levels can also be biased to favornormal operation. Subsequently,signalsare split to record low- and high-frequencycomponentsseparately.
l_w-frequencyor quasi steady-statecomponentsare generatedby low-passfilterstl_ateffectivelyeliminatefrequencycomponentsabove 0.4 Hz (3 decibellevel). Signalsare sampledby a sequentialanalog-to-digitalconverteronboard the aircraftand stored as digitalwords on a tape recorderlocatedin the crew compartment.
lligh-frequencyor dynamiccomponentsare generatedby high-passfilterstlmt effectivelyeliminatefrequencycomponentsbelow 0.4 Hz (3-decibellevel).Signalsare furtheramplifiedbefore passingthroughvoltage-controlledoscillators,multiplexed,and recordedon the same tape recorder (differenttracks)used to record the low-frequencycomponents. With 21 continuousbandwidth (CBI_)signalsmultiplexedper track, four trackswere requiredtorecord the 80 high-frequencysignalsfrom both instrumentedinlets.
I)uringthe zero and calibratesteps, all filters are bypassed,and thetotal signal (minusthe referencepressure)is recordedboth as pulse-codemodulation (PCM)and CBW parameters. Calibrationsequenceis operate-calibrate-zero-operate.Transitionbetween the zero and operatepositionsiml_sesa step change in pressure differentialacross the transducerdiaphragm.This has no essentialimpact on the low-frequencycomponentbecauseof thelow-passfilter. However, the high-passfilterrespondsto thisstep changeand res_altsin a damping characteristicwith a period of approximatelyi0 sec-onds. As a consequence,dynamic data are invalidduring this time, and datarecords for detaileddynamicanalysisare selectedto circumventany problems.Typicaloutput signals, includingcalibrationcycles, are illustratedinfigures106 and 107.
102
Parameter identification
(scale in counts)
PT = (CTSOP-CTSZ) CAL + 6002 CAL = (6002-6631)/(CTSOP-CTSCAL)
"' '" 7"'_oPE.CTSOP I , ,I CTSCAL RATE
_ _3_6 _, ....i_---_I/t-l, I 633CTSZF_ , -r I I ....
*.o _ .... I,..... ,,,[u, ..... !1.... ,I.... L_,..,u.JJ,,,I o.o
6327 ' I I 6332
...."......' _........_ L....I ....... ' I : : I'[ .... : _: ": ' r _ I i i1 L.... i J,., .......... , ..... ..;-.j............ ,., ' '7
, , , "'"_• I I , ' --76328 ...............,_ i_ __ I: CALIBRATE .-,.._I'_I I
6333I I''''......,--1 ] ............... P--i mr......,.o ,.................. L......... _........... ,., ,........................ _............
6329 , i t 6334 .
................_ r _ Zili,i III111111 iiiiiiiii iiii1111 iii iiii Iii i iii l.i I
-Ittl ....i [1t6330 ', t 6335 I I I
I I Ii,O __JIIh,LLIII _'_tlT nit Itl O,I IIIllUll Illllltnl itQmatllln iJll*_
• leO _o _o _oo _0 too loo _o _eo _€
6336
'_-' _ 1 -..:r,,.. ......,_.....r,......L......... I,L. i t .......O.O '' I' IIII ill ill , , , 11_ l, , ,, , , , , , 11 , , , , , , , , _.O ' ' I I ' I 'I '
-LI .....iSi, I u E I END OF DATA RECORD I
6337 ; , , 6342 .................,....._............ -4 i .....
O ............ _ ...... _1 _ ..... "111 , , 0.0 , .................. _ ..... _ I'-| L&
..... I "gt t'4L't,1''6330........... ,, ,, 63_3 ..... 'I.I lima InlllII I'_" hie O.l IIII n l.IJJ lIlIIlnIeIJII.U I1_#
..... 13ltt ,].....I _ I f I I
6339 - "............. I.... _.__.F.... 634-_ ............. [" , _._L/-..... .....
_" ,.,.....................I.....,li....I................ ,., ....................!.....' _* ,_,,,,,
°LAtt !J .... , , ,.... I l" .... I ,',"I I "- j-I,I I._. .ii iiiiiii ii_ ii I.I
• Ill Ill li,OI '_ll log IOl Ill loll i00 Ill
SAP.1PLES SAMPLES
Figure 106. In-flight calibration cycle, total pressuresat aerodynamic interface plane, flight 1-5, 64 sps.
163
_" F--- - -_ ................=--- - ' -:_-- _----- --- ---_ - - --=- ........ ............-------- _ " =O',: - -= : --- = ..... - - -- --- =IL_+4V FS disc
'- _ --- I _--- ......-_- --_- -: :-output (62.5 Hz)6 8 .... -- " -1 sec -- _:-- 32 --- :-_--_:----= ------° _= L=_--_ .--=_..... -_ _ ! - - :-
erate .-Zer Cal __- Ol_e -- " ---_-._- --- ---= -_=- ----=:- ___
V=i o fi, --igh-pass N (High-pass --- .... -i ..... =L_ L- -- -- ___--_--=:__=:-Z-__--- ! - - ---_- _ NO valid CBW data during filter __:
] .... _ .... . ........... =-.................. I _ . --=. _ stabilization time of 6-I0 seconds _---_ I --_=-' .... _= =_T_._-_- .......- .... -- 3 _ _-= ....._ i_----_...... _ -:--_- ---- ___.._-1-._ -=- -1:_ - _ _. - __..... _ _. ___.... _ . =--.:_
...... I_- _ -- _-- _=FT_-__-=- -...... ,.. I _.-z.._.-_-_-_--. -_ -_- .... _- _-- _ ............
I-_6329------- - .... - ...... - -- = =---_- - -_-==_ - ------ ---= ..... -= .... - ........- __. _=__- -_ ..... ___--.... _-_ -- ... __.___----_.: _--:=___
_- [ . .' _- _-- ._-'-=-_.-_-_=_ ==-_..... _ ...... I - _- _-, _ _ -_ -. _ - _.-
+- -- _ I _-- --- "...... - _=-_:_=-'--...... =---_ ....= " - _..................... - - i ...... L ....... -- =...........................
_ ............. _ .....
........_ _-_ ......... =T ..... _ i -- " .......... i r£-.... ____ . L ............ -- 4- ---- =--
- - - -'-_;- ] __Determined from PCM data-_--_ _ ........... --_--- ...... _-_--.... _.-9:- . CTSZ_H ....... _ --_ -....... - --_...... _ - ! - -
I
I m
6330 =_ -- Ap- _ CTSCAL ................ _ _ ....-- _' -=_==_-=-=-----_-_:-_--T_ _.L............................ _ CTSMEAN- __.....=....-_.=:_.... ...-.................................. _. _=-Lq ..... !
_- I , -_I ..... _ I _pJ- + PPCM.__-----=_--_...... - _..... _ _ ._-_. PDYN: (CTS-CTSMEAN) (CTSZ-CTSCAL)__
......... ..__Spur_ous ZOC valve operation ....... _ I_-_,=__- _: - , _ ._ _...... .j=-_ ---.. _ '..... i _ _'==.... - ..... -"....... -_---
__ ' __-. _ _.... , , _ _ ,i--6331 ..... .......". _ ..... _ _ l t_=_ ........ __=-__-=____==__ r...... ]_._.-:-d..-.---! ..... -..... _-
- "_ ........ 4- ...... :---= ....... ] F I _
_Ap.p=_rox t_ime l :27: 14, fl-t- -- I ,- ---_.. _ - _ .... ] __ - ..... --_ ......... _ -- _
_ -.-- =-_,,_ ____. ' " - =----2-.-- -
Figure 107. Repmesentative discriminator output, AIP total-p_essumeCBWdata, £11ght 1-10.
In stmm_ry,particularemphasisduring the flight-testprogramwasplacedon obtainingaccurate,total-pressuresurveysat the inlet/engineAIP.As described,these featuresincludedthe following:
(1) Differentialtransducerswith a floatingreferencepressure
(2) Accuratelymeasuredreferenceand calibratepressures
(3) Automaticin-flightcalibrationprocedures
(4) Signal conditioningto obtain good resolutionof both high- andIow-frequencycomponentsfrom the same transducer
F1ight-testexperiencehas shown that, as one might expect with all thepneumaticand electricalconnections,the systemrequired considerablemainte-nance. However, it is believedthat system accuracyapproachedlevelsavailablewith a well-constructedwind-tunnelprogram,and this seems to beconfirmedby comparisonsbetweenwind-tunneland flight-testresults.
TEST RESULTS
A total of 5 flight-testhours were dedicatedto exploringSMCS vaneeffectson inlet/enginecharacteristics. Flight conditionsand operatingvariablesare summarizedin table IX.
Several computedparametersfrom the 40 total-pressuremeasurementsat theinlet/engineAIP are used to present test results. Recovery (PTI/PT0)is theaverageof the 40 low-responsesignalsreferencedto free-stream,totalpressure.
Several distortionindexesare used and were computedby digitaltech-niques. High-responsesignalswere filteredto 62.5Hz and sampledat 360
samplesper second to representa one-per-revolutionengine-frequencyresponse• (signalsfully attenuatedat 125 Hz). Circumferential(IDC)and radial (IDR)
distortioncomponentswere computedfor each ring (eachgroup of eightAIPpressuresat the same radii) and combinedmathematicallyto form a fanstall-marginindex (IDL). This latter index is normalizedto stall-marginallocations,and thus, values of unity computedfrom the high-responsesignalsidentifydistortionlevels approachingdesign limits. Inlet distor-tion, (I_MAX- PT_,IIN)/PTAVG,was also computed. Record lengthswere in therange between 5 and 20 seconds. Scans resultingin maximumvalues of stall-margin index are generallyused to identifytrends.
165
'I'ABI,IiIX. - SMCSVANEEFFECTSON INLET/ENGINECHARACTERISTICS,FLIGHTTEST INVESTIGATION
Dedicated AlPflt timea instrumentation
Flt no. (min) Objectives/accomplishments status
2-19 65 Initialflight-testinves- Referencepressuretigationswith oscillating transducermal-vanes at 1 Hz, vane angle function,no con-±i0 degrees, u = 3-8 degrees, version to= 0-3 degrees, initial engineeringunits
throttle transients
2-33 45 Static vane deflectionangles ZOC valves mal-in 5 degree increments, functionedon± full scale, u = i, 3, no. 1 eng, data5 degrees, 8 = 0-4 degrees, reducedonly forno. 2 throttletransients no. 2 engine.
2-36 60 Repeat of flight 2-33 Leaks in refer-
- ence systempre-cludedAIP datareduction.
2-37 55 Repeat of flight 2-36 Satisfactory
2-38 60 Max rate throttletransients Satisfactorywith vane deflected20 degreescombinedwith nose left
sideslip -
2-42 20 Conductpushoversand pullups Satisfactorywith vane deflected20 degrees
aTotal _5 hours
166
Engine-face,total-pressurecontoursare used to illustratevariations in
distortionpatterns. High-pressureregions (pressureshigher than average)are shadedon the contourplots. Low-pressureregions (pressureslower thanaverage) are unshaded. The magnitudeof the differenceabove or below theaveragepressurelevel is definedby the number on the contour. The annularregion is formedby protrusionof the enginebullet nose at the AIP.
Envelopesencompassingall combinationsof circumferentialand radial
distortioncomponentsare used to identifythe magnitudeof dynamidactivityand to comparethem to design goals. Two types of time historieshave also
been found useful in documentinginlet turbulencecharacteristics.Analogstrip chartsof the high-responseAIP signalshelp to identifywake ingestionduring transientmaneuvers. Digitaltime historieshelp to define transientconditionsbased on computedparameters.
OscillatingVanes
Initialtests with the SMCS operativewere conductedwith the vanes
deflectedsymmetricallyat nominalangles of +I0 degreesand then oscillated+i0 degreesabout that mean at a frequencyof 1 Hz. Vane deflectionanglethus oscillatedbetween 0 and _+20degrees;the latterrepresentingfull-scaledeflection. Note that with these procedures,the vane is being used as anexcitorand inducessome discomfortto the crew members. With the vane
oscillating,aircraftmaneuversinvolvingcombinedangles of attack and side-slip were performed.
Time historiesof aircraftattitude and vane deflectionangles recordedduring a 30-minutesegmentof flight 2-19 are shown schematicallyin figure 108duringoperationat Mach 0.83. Angle of attackwas varied between 3 and 8
degrees; sideslipangle was varied between 0 and 3 degrees. Tests were con-ducted at positive sideslipangles only to ingestthe wake in the instrumentednacelle.
Flight times where increaseddynamicactivity (identifiedfrom stripcharts of the high-responseAIP instrumentation)could be definitelyattributed
- to SMCS vane deflectionare indicated. Increaseddynamicactivitywas alsonoted at other times; however,effectsof sideslipand vane deflectionanglescould not be definitelyseparated. Ingestionis generallyrestrictedtopositivevane deflectionangles (leadingedge up) greaterthan i0 degreesincombinationwith aircraftsideslipoperation.
One major advantageof oscillatingthe vanes is to producerecognizablewake ingestioneffectsin the data. The 1 Hz vane oscillationbetween 1.5 and
18.5 degreesduring a sideslipmaneuver is shown in figure 109. Time histories
167
Leading
edge up L_.Vaneosc'illation at l Hz20 .::.:;.:r.:: :.:.".-,..:.:: .... :.,_--_-....:::.r:-!::=========================l": ....:-: : . ).:! = ",.: :'".... :..... t'--":'::;' _:(.':-_-::_:"..:-T-7_T ":"T-.-T--F.- " ":-::---:._.-:T:..':_'-".'T,7:" "_'7 • • , ;. _ ":: -" ::"=.-:i-
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u mmO " : 4 : " 8 ....t:; : ]2: .... ... 16 .-i 20 . • 24 28 i.... :..-::-;:-::: - .-.-.:- :-. -:::i ...... " " : :";--:!_-!7: _..:::iE _ " : :"' : " " : " " " Flight time - minutes_I_ _ .... " ............... -- .................. ...... :-_ r': "T-- °e- _) : "" NO. 1 "_- I • : ! • " : .... :" : :':.":::";"::: :.: • ........>'_ _I_. = ....... ,
_ ..... inlet:2. : :___ . _i:,i ::--i.-::: ;:-: Increased A-IP dyr)activity_:-:.-----'_ • .i ] .:i.i due SMCSvane deflection"_ u " No.•- _ ' inlet ............................... "..... ;
__ _ l:igure 108. - AIS/SMCS investigation - £1ight 2-19, Math 0.83.
168
0 ,,, -
x3029RHSMC 0 _ ]Vane - _- l Hz-
I
- SMCS vane deflection angle, clegrees
--20LI I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I !
5 i . |
- i Sideslip angl- .. -.'T '.I . _,degrees.
MlOl2 0Beta '_
-5 I l I If I If I I II II I II I I I III I II I II Ill IIII I II IIII II
2400 , . _ ,- AIP PcMdata
Counts
00:53"40 Flight time- seconds2000 I I I I I I I I I I I I I I I I ! 11 I I I I I I I I I 11 I I I I I I I 11 I I I I I I
O 2 4 6 8 lO
I .... ._-_:..--.._._'_L -_ '__,l .... W-data l.."_:"-:- ......Representative AIP C-B I .......... ; .............. '.......o !L ...... ._..'_'--1 ....
= D6338
: _ High pressure> " F.,_4-J
m D6339
' • ' ......:i_ _ Low pressure: _ , -,, _ i ",._-
Figure 109. - AIS/S_S investigation flight 2-19, Mach 0.83/16 000 feet,
= 3 degrees,vortex ingestion in no. 2 inlet during sideslip operation,PLA = max.
169
of both low- and high-responseAIP signals are shown duringan increase insideslipfrom 2.2 to 2.7 degrees. Wake ingestionis evidentthroughoutthetransient,and turbulencelevels increasesignificantlyat the higher sideslipangle. Although the individualplots are not preciselyaligned in time, lowpressuresgenerallycoincidewith positivevane deflectionangles. Similarresultswere obtainedwith the no. 1 inlet. (Seefigure ii0.) Smalleramplitudesare partiallyattributedto reduced engine airflow indicatedbylower fan correctedspeeds.
I_gineairflowdemand has a significantinfluenceon peak-to-peakampli-tudes. RepresentativeAIP pressuresare shown in figure iii during a throttleburst from IDLE to INTERMEDIATE(maximumairflow)power settings. Peak-to-peak amplitudesat INTERMEDIATEpower are greaterby a factor of 1.7 than theamplitudesrecordedduring IDLE power. Again, the lower pressuresin thecycle appear to coincidewith maximum,positive,vane-deflectionangles.
Static Vane Deflections
Comparisonof resultsfrom tests describedpreviouslyindicatedthatresTJltssimilarto those obtainedwith oscillatingvanes could be obtainedwith the vanes positionedat a constantdeflectionangle. Crew members found
these proceduresless objectionable,and their work load was slightlyreduced.All subsequentflight tests exploringSMCS vane effectson inlet/enginecharac-teristicsduringmaneuversand enginethrottletransientswere conductedinthis manner.
Maneuvers.-Tests conditionsinvestigatedduring flight 2-33 (withSMCSvanes deflected+20 degrees)are summarizedin figure 112. Tests were con-ducted at Mach 0.85 and includedmaneuversat combinedangles of attack and
sideslipwhere the maximum angle of attack was 6 degrees and maximum sideslipangle was 4 degrees. RepresentativeAIP high-frequency,total-pressuresignalsare shown duringmaneuverswhere vane vortex/wakeingestionwas measured inthe no. 2 inlet. Faulty ZOC operationprecludedanalyses of data recorded in J
the no. 1 inlet. A strip-charttrace,recordedduringnormal cruise attitudes,is shown for reference. Aircraft attitudesresultingin ingestionduringflight test generallyagree with those recordedpreviouslyduring wind-tunneltests.
Steady-statetotal-pressurerecovery and distortionparametersare shownin figures113 and 114 as functionsof sideslipangle. Angle of attack is used
as the independentvariable in figure 113 with the vanes deflected 20 degrees.The largestdefect in recovery (approximately0.06) was measured at 6 degreesangle of attack and approximately1.75 degrees of sideslip. Effects of the
wake diminish as sideslipis further increased. At approximately3 degrees
170
100 ------..------r-----,.--------r------,Percent fan speed
Mll12NF
f-
-iSidesl ip angle, degrees
,i-
i="
i- ,
i- t·i- .'i-
,f-, I I I I I I r I r r r r I , , r , r , , , I I I I I I I I I I I I , , J I I I I I I I I
- 20 L1..LLLLLLLLJ-L.L.J....l...1...J...L...l...1......L.J......L.J....l..J...l..J...l..J...l..J...J...L..J...L..J...L..J...L..J...L-1-L....I....l.....I....l....J.......J.:o
5
OLLLLLLLLLLLL.L.L.L.L.L.L.L.L.L...L..L...L..L...L..L...L..J-l...J-l...J-l...J-l..~~~..............................~
20 -----.........-----.....,...----~-----r_---_,
-52800 ------,,------;,.------r-----,------;
Ml012Beta 0
X3029RHSMCVane 0I------~------+----..,---I-----_t_----___t
D6288Counts
AlP PCM signal. I
40 CTS at 0.01 0.4 psi
Figure 110. - AIS/SMCS investigation, flight 2-19, O! = 3 degrees,vortex ingestion in no. 1 inlet during sideslip operation.
171
100 ,..-----.----r---"l:::::===---;t;r-----'
INT
Mll13NF
Percent fan speed
tIDLE
.- Sides 1i pang 1e, degrees::::J
I--
I--
l- tl-
I--
tu-u.-LL.LlJ I I I'll I I I II I I I I I I I I I I 1 II I I 111111111
OLU-l..LLLLLl--L.LLL~__LL.L.LL.LL-L.L.L.L.JL.LL-L.L.L.L.J_.L.L~L..L.l_L.L.l.._l_LL..L..J
20,.------r-----,-----,------,-------,
- 20 LU-l...l....u....l..L.J-I...L.L.IL..J....l-l-L..J...J....L...I..~L_1_l_l...J......L...L...L_L_...............................-'-'-..........'__'_'_...I....I........
5
-52400 r------,-------.--,--..-----:T~--_;:.V""--r___,
X3029RHSMC 0 I-------+----:.---+------+-----+-------tVane
Ml012 0Beta
ps i
Flight time
2 4 6 8 10
• .• -- ... -'. - ·.A. jp -.c"sw- ;i~n.~.l·~. -- ~'. -:1.---... ~ f.-. ---.1.".;; . . . . . 'I --- i ... JJI _.•~ ... ~
[ 06338 -'\C~'(-t~~"'''li~ti"J~~r-~~\''r. ~I.· -- ft11;VWI.~I'~~~----: .Low pores sure -" ---,----1 . 7X _. . -- '_'_'_1 .'--;
!D6339)IR~~~A~~~,:Ji~I~~~ .~~~_. !~\~I_'III~...~.. ~--_.::. ...~.:.f.:" -.~_.__LH, gh pressure . .' I Y: "; .._l~\·
Figure 111. - AIS/SMCS invest;igation, night 2-19,a = 3 degrees,vortex ingestion in no. 2 inlet
Juring IDLE-INT throttle transient.
D6338Counts
172
Fit 2-33 test matrixEnvelope of SMC vane 10 Angle of attack,
wake ingestion from degrees • Normal turbulence Normal AIP dynamic activity
wind tunnel data 8 O Increased turbulence (No vortex ingestion)
....Throttle transients -i_i_-- _ _:i '7-J_'2-_ :_- "--: - -i-"--_:--_
• _ --- ....... :- :
Mo re " _6 _0.3 psi -" - -_ "
4 6347 -_-....:_:-..:_-:_---:-:--._--_-_-_4.-----
...... •..: _: ..... - . -
"!"i!i!iiii:_":.... 20 : 26 : 44
0 oz= 3° _= 0°
-8 -6 -4 -2 2 4 6 8
-2" Sideslip angle, degreest
AIP dynamic activity with vortex ingestion
_:-=--'--= 2.3 psi .... _ .... 2.4 psi ................ :----:'_........... 2.45 psi :_--_
A,P probe _--.......=-_=__.:_"_-__:__L"_-=_.=_':;_-.._;--:_-_--_--:.:::_:"_.............._: _-_,-_......_ _- _--i_ii_._____-_-i._-_.iL----_i i:_:_!--__._.
19:58:00 20:28:I0 20:43:00
e = 5"5 ° _ = 2° _ = 3° _= 4° OZ= 1° /_= 4°
Figure 112• Flight 2-33 test matrix and AlP signals during8ach 0.85 operationwith SMCS vanes deflected20 degrees,
no 2 inlet,RB = 7 degrees,RC = 5 degrees
INo.2 INLETIM= 0.85, alpha = variable, beta = variable
RB = 7 deg, RC = 5 degSMCV = 20'deg
l.O0 Recove ry
0. .......:!:
0.92 .....i i.1i.:......:i. /;!i_.:__..!:: ......i.i:.i. :..
•.. i.--i-::!.::i.-ii-:_!Sym Alpha WIR .i_i...]..i!ii-ii:i:i:.-::
•' : :--:: "i_":i O I deg 297 pps ::'._.::. " : ::: :..---"..............-:.-;.A 3 deg 323 pps .!" ::.-.!'.:::i:[.i:::
• : :. :.:':.:.-:: :; {'1 6 deg 348 pps .i. : .i ...i. iJ.i.:
0.8 ................... ::; -:: _-:.Stal.. 1-margin indexl ._ -:-i!..::i:fi.::i.::.'..:. .; . : . .: : '.;::;.- .L':::: ":'..:::-.;:L-';'.:::_:::: i:::: ".; ' ;.'.*._ . :; :_ " " .
IDL .... " .......... ' ...... :.........J.................... :...... :........ : '-.','.' :..': ....
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. ' . ;. --. ...... -_" "i ................. ;_-% .... _" ;__T.Z_............... , .......
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•; .... ".:-::::::.';:::.'--'::'::-: -::;: D stort on .,-: ..; .... -;.,:--::::Ti:.:'..7.--T
0,2 .! .:_. ! :.L:i::<:i:i.::::.:.:.;: .::{PTMAX_PTMIN]! :'! : ....... : ::.:-1---::."
I DT :. : .:..'I:. "; ;I. :.'::::.: ::- :::::::::::::::::::::::::::::::::::::: F. : _ :T.:';_ ._:.::.'..
---A'-"'_.........:_. . TT-,-_......".....;'-'+.........._,.... -t---_._'--AL-..T.......*......................."-- -
:::_ ; : _- : : : : : _.;:. ; - ........ : A: :: :::::::-
.---:.+;+--:--.:---=_ i._ +:-::.:H-_ _-::_:_::__:::_::::.?'_::.__?:+:__:_:_ _::: _
- 0 ! 2 3 4 5Indicated angle of sides]ip, beta, degrees
Figure i13. No. 2 inlet, effect of sideslip angle on steady-state inlet
characteristics at SMCV = 20 degrees, flight 2-33.
[74
M= 0.85, alpha = 3 deg, beta = variable
RB = 7 deg, RC = 5 degl.O0 . Recover,'
L........._...-.%. _Z_.:.:Z_:£1ZZ. Z5£Z.!'.-Z.:'-_-
o.96: !• i. ' : : ".__i" "'.:[7:_'.':::':x:..:x:t'.:T_x::t_":. _ ':;.. ;. .! " ' : .:i ..... ,_......
PT1 . ...: .:r.:__l_ i..q.::_:::::b_:}::::!.::::Fr:_::::':.:t..::': 4 : 'i i:V :i:! " ::-
.:: .;.::L:__:L_:l.:j:_Ti:£::i_L?i£_£{. ,::.: ......... r:'::..._ . ::::..:. :::': :.:-:_ .:-:". ,- :::.:- :" :;t:- .-"-: .... :'--,:"%::: "- ::::::::::::::::::::::::::::::: "- :- :!:'-':"::'" - : _'::::::-
0.92 ..--_-.......... -.....: '!....."...............•i'..'.::. '. ....:.... ._. :... :. .....:: ..... SYM SMCV WIR .!::::i::...... _. • . .
;-N_t.; + .:i:> !:.:_;_.._:::i?_::>:.i :-: ,: • '-:-.::! ::.i_"_-_:! _ 0 316 pps-:---i_i"
•" : : i'. ":".'t:...!::.:_'::::'."::-.. '. i l d :: _.... : :"___-.._[. :.!_::{__.:k_:____.-.._._ _ 3.3 eg 316 pps ._-:_-.::.i!_:--:-: -" ::. :-__'_-_'_!::! g ,20 deg 323 Dps :'- :": ::.- :-. ::::::::::::::::::::::::::::::::::: . :_:l:.:_:::.!_k ":::: :.: ::.: :i ,. ..'._
............... I ..... _.... _ "" _....... _........................... _..... _.......• : " [ : ::'":::-:::: I"':"' i::'.:::|:-';" .:.t::-::::-:l -':". I" .': "; "" :: .:.'.. :.: : : : "
0.8 ." _ ............:..........r,-_-_...........:......._:_:._ ...........: . ....-.:;::_.--:- -t:::!::-it!N--!:iStall-margin index :::q.;::":i.:;::-i -.--.-
IDL ::_.L. .. :i:-'r-.:.i-'-.i:::'::::'{::: ::]': :::: :i.: :::. ].:'...... l ...... I: .............. 1:7::............. J......... , ....................... t-..........
-- .--..--7.: -_-'--_.--_; _'K "t:-- --_.- . .._: .... !.... ;- - -' L :-_,--7"-;"I---. _--:v"rt ....... T---'.'--"7..-_ ..... :--':C-,.---T
0.4 :.:.......:....": ........".........' ...."="_' ..... i " _......i.........
ii!!ii:!Ni))i:16%;i'-.--!-iii;iHSii iiiii,iiiiiiiii!!ii ;: ii;,iiiii :0 ._::i: :.} ,:::::!:::.':.: _:.:.:.-:::k:_."_!:!:.:.:._.:q:_!::-:i.::.-- i,:::'-:':!:.:: ::::::::::::::::::::::
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-".:. "._,':::::.:'i::xi:-::_'::::':::}::::::.:.,t':.::,:r_:*,-_n::::it_'::.:x'::: ;:. ::::::::::::::::::::::: _.::":_.:':". _:_._.A_;:_r%:_:_::_:'_:::-£._T:t.N_._[7_r:_.-_.:_:_:_::N..:.°" :.L.:_:," _.:" _.:::._t:t:: _. ::...... ,......... t .... -..........,.......!.......... _...... -+--:r=.-" I- ........................... • ................
._-" .: _" ;' "'_....... :" ',7-"'....... ! ......... i .... T............ .",, ....... ,...... ' ..... ,."............ -..'.::.=======================================================================: ::... :'.:'._1::_'.:;:1:" ;; ..... "
.................. --I .... I ......... _.... _ .... _............... _.., ....... _.... , ........... | ....... . .................... _........ ' ........ _........ I ........ _"....................... _........ l .............
0.2 :: ::.:.i}::}} :. .:.::_.:::ib:::::_:!.:::!:: Distortion .:-!:. !: :_: ;:::-!-::._::-!:.::.L-::-" ::::.Z :::::::::::::::::::::::::: _::ZZ---L--!-:Z[LL:::Z-_F£f.:-
. : .-:..:]: .:.:..:::::::::::::::::::::::::::::[PTMAX-PTMINI-..:........:.-. ii::.-:!.::.i:!:qi:.!I DT -:iii ::i i_i[ii!i:.|!!i_-i:_-i::i:i:_i!ii:i[_iiii L PTAVG /::.::!i:::i:::: i -:ki:i;:i-!i:;".::.i::.J
.._£._i....:.:J_!=:_::::i:::i.k.:::!_Z.-_.:::L::_::_:_:..... _i£_i..::_i--..-.-:_.:_:_ :: ::-L:. _. .] ;:.:_._IL:_:_::_b:.'_i:_:.k-_::.!-_:.iL_:_;_:_N::.:-_!:._:..-!_-...:::_: _;II:.-L-: i:._i:-
-.- O. 1 . :. :i ::: :..1:.i_:.:.:-:.!.ii:_:!:!:.!!:::!::::!-.iiN-{ii::__._.'.::,iii. " :: (_Z_L .:i.:: :..-._.::.:::- .::- ._ :.h:::L-::,:::_i-"Zx:_:":7_:::.]..,,n:.:_•. _ki: 8;.:-i:_:.t__:._._
- 0 l 2 3 4 5Indicated angle of sideslip, beta, degrees
F]gure 114. No. 2 inlet, effect of sideslipangle on steady-stateinletcharacteristics at various SMC vane angles, flight 2-33.
175
sideslipwake effectsare no longerpresent in the inboard inlet. At lowerangles of attack,sideslipangles greaterthan 3 degrees are requiredbeforewake ingestionbecomes evident.
Effectof vane deflectionangle is shown in figure 114 during operationat3 degreesangle of attack. Pressurerecoveryvalues from this figure indicatea similarloss during sideslipoperationfor both the 13- and 20-degreevanedeflections. No losses in recoveryare evidentwith the vane held in a neutralpos ition.
No definitivetrends are evident in the steady-statedistortionindexes.
floweret,with the measured increase in turbulencepreviouslyshown in figure 112,this is sometimesindicativeof an in-phasecomponentpresent in the flow asshown in the time historiesof total-pressurerecovery (figure115) where SMCSwake ingestionwas most evident. Time historiesof stall-marginindex are alsoshown. Similartraces during operationat normal attitudeswith no vane deflec-tion are shown for comparison.
Resultingenvelopesof circumferentialand radial distortioncomponentsare s}_wn in figure 116. Total pressurecontoursrepresentingmaximumvaluesof stall-marginindex are shown in figure 117. Maximum dynamiceffectscorre-spond to operatingconditionsresultingin lowest total-pressurerecoveryat6 degreesangle-of-attackand 2 degreesof sideslip. Maximum stall-marginindexwas 0.74, and the associatedtotal-pressurecontour shows a well-developed,low-pressureregion in the hub.
An extensive series of tests with the SMCS vanes was also conducted duringflight 2-37. Performance and distortion characteristics are summarized for the
11o.2 inlet at normal aircraft attitudes (_ = 2.7 degrees, _ = 0 degrees) infigure I18 for the complete range of SMCS vane deflection angles from 20 to
-20 degrees. No indications of wake ingestion are evident.
A seriesof tests investigatingSMCS vane effectsduring sideslipopera-tLon were conductedover a range of angle of attacks. Resultsare generallyshown for both inlets,and in each case, resultsare shown with the vane at aneutral positionto help separatevane deflectioneffectsfrom sideslipeffects.Re_Its at a nominal1 degree angle-of-attackare shown in figures119 through123, 2.5 degreesangle-of-attackin figures124 through126, and 6.0 degrees _-angle-of-attackin figures127 through132.
Generally,these figures illustratethat the higher the angle-of-attack,the less sidesliprequiredfor the wake to enter the inlet and affect thedistortionpatterns. Data recorded during tests at 6 degreesangle of attackare also used to illustrateadditionaldata analysistechniquesbesides thedigitaltime historiesof quasi steady-statemeasured and computedparameters.
170
, . , .
a= 5.5° {3 = 2° SMC vane =+20°
PT1/PT~ ,0613:40.8 11 '111 ~ I
Fi Itered to 62.5 Hzsampled at 360 sps
a= 3° {3= 0° SMC vane = 0°
PTi IPT:,: l... 0 0 0 0 0[J"""", I, " , '" "I
20: I 9: 57 I .0 2.0Time - seconds'
1. 0 ...----,.-- ---r-----r----,
IDL ~J\'1',~t~':~:\V!f~, \'i/\>-'~tJ.l.I, \Wt,\I,r'lf\'0.0
19: 57: 49 I .0 2.0Time - seconds
a= 3° f3 = 4° SMC vane =+20°
PTl/PT:':ttt=u
2.01.0
1. 0 r----,---;r:- dIDL
o.0 ..........-W-L-Iu....L.J.-L~U-LLLLl..I....L :: L1-1..L1-1.
20:43:03
Time - seconds Time - seconds
Figure 115. - Time histories of total-pressure recovery and stall-margin index, Mach 0.85,with-~~S vane deflected 0 and 20 degrees; no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.
o.og.oo 0.02 0.0" 0.06 O.OB 0.10 0.12 D....IDC
IDC IDC
Figure 116. - Dynamic circumferential and radial distortion components, Mach 0.85,with SMCS vane deflected 0 and 20 dep,rcc-~~no. 2 inlet, RB = 7 degrees, flight 2-33.
, ,
Pressure above PTI
Filtered to 62.5 Hz in shaded area
sampled at 360 sps Pressure below PTImax IDL scans in unshaded area
,number onPTI - PTI [I+ _ contour ) _ ]
Use + if contour in shaded area
Use - if contour in u_shaded area= 3°
SHC vane = 20°
Figure 117. - Dynamic total pressure contours during Mach 0.85 operation with SMCS
vane deflected deflected 0 and 20 de_rees, no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.
Recovery Flight 2-37
1.00 ?._i_i::i::.::.-.: ::.:: _-i::.:!::!-ii ,!:: . " -:-: . : • ".................................. : • : T.. ::...: ,, • i ....
0.98 ?:];::?i:i-_"-i!i_i_i_iiii;:i!iT-; I NO. 2 INLET I i!71:!]!:. !i=:::7:_i:ii=:i;i:!ii_i: ;
:=LJ--.L__:-_ RB = 7 deg_ RC = 5 deg _-............' _ . "I. : : i_.;..;...PT1 _:-:.. -- -.PTO .... .-! _. M = 0.85, alpha = 2.7 deg, beta =Odeg. -.;: : '...
_:4_:..J L_::_2..o " :.: 2........ L •-: : 1".: L' _- :&:-.Li_.£ .'_- :__.__:.i:.__'_ _'..._:41_...- :._.-L_LI: --' "'_ """ .'.:L_. :.:..L: _L_I
96 ............................................• . . 2 .0 -, - " ...... I .- : . -: :: -, " : -;" -; ." ;: :::,::;: ::." ; . -. 'L'.:. '; "
• ,q .... :-j.= : : " : _" ; , " :;: =I " .:; ; " :..; '. :I. ; . " :: ;:: : ; •
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::: :":1 .. :-.:I:':= ",L=:'::'._::':' ._:.:',.: .... 1=:*. ; ;. :'. L :: : "_ "" :1
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I DL 0 4 ""::::::" """ _:; :" _': .... ":!':: "::::_' .... i " ..... _-; .... ---. ,.. ; ._ ' ..- _ .... =o_ . . _ - " _ --_ .
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.-:_'-% ;:i.=_4_-__;ii_P_;ii.%__:_ :__:_t_:_ i%_._L %:_4i_"
". ':::.'. =:'-:-.:'t:=. ::::::::::::::::::::::::: u::.::".;::'::.'::::: : ::i-:F:_.-T::::I.::2-.'_TT-:T:IDT 0.1 .......:.... :: .,._::::_::"'':":-"::;:::=:::::_:.:...... _..:: .......... _. : ===================-===========t=====,, .......
::::h :.:ii:.]i::::!i."_ ...... :-:;:::.='::'v ::::!!::':::::l':::]:7"h::'!:!:!:: .':i !": i' _ . :." :............................. 'LLI:-_:--2. LZI...L.L: "_-i2:=J%__.L." "' -- • " L =.= " __.-'_ .... ;_. "......... _: -F:J.............,.... _ .......T ......-.--_.._ ...... - -_4: ::: __::!_:=:f_ =::..;m"_..::.A.: !:,,_::::_!.:G).I_.:_::.::.-._:: "
...... =- _ ..... l ........ _v ........... ----_,,i. _ . _ -- " ....: " ':"_:"::-_ :::_:--::=i:i::::i.i:!!i--.:i:" i"i:t-::_'::=i::!-!.:? -:::i!-'
: i-::::f:!'iir-:;':-.-.!:::!-::T:_:_---, ....................... _....... t:-_--:::: .....:i ................... ! ................................. " ......
-20 -10 0 10 20
Structural mode control vane deflection angle, deg-rees
Figure 118. - Effects of SMCvane position on no. 2 inlet steady-state anddynamic characteristics, Mach= 0.85 and alpha = 2.7 degrees, flight 2-37.
[8O
Mo (PEG) I
0-
-" 0.5 JII LI[1111I_ 11111 L.LLLL -2 I_I_.L.L.L.J
ALTITUDE (DEG)(K FEET)
o IILAI''' ,,,_,_i,, ,,llJllll k_,,,,,,J ,_,,i,1 .... ,, ,, .5 IIItllll_JllJlllllJqlllJL_
'% I I t I I 25 'E -- I STRUC'_JRAL HOOE CONTROL VANE _SI_ ....
-_0 J.tl tl[ll ll_ 1_ U!l[llll tlllllltl !ILLL2LJ. -2_ 1 JJJ.J._LLLI lJ0 _.0 80 120 0 _0 80 120
_ 20: k9:32 SECONOS SECONDS
I No.l inletI INo.2inletIRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg
_ i __ _ LO --_ ..... , STALL_RGIN INDEX "-- 70 ]1 f t 111[ I III 0 I III , . _.... _ ,
0 Itl[ ]1111 IIII Iiiiiiii I I II I 1
°" r-i ]FAN DISCHARGEHACHNO. SIGNAL GNAL
(AFT CiNTROL)
/
iitiil[i itll_l:_ ::ii_111 I.i]ll_ll II!ll;llI_[illlll L 0 F_l_'l_ll/t:_ll_ lllrillll: :1, _:.: t_--.--._._-L .... _L_I
0 gO 80 120 0 _.0 SO 120
SECONDS SECONDS
Figure I19. - Effects of sideslip angle on steady-state inlet characteristicswith SMCV= 0 degrees and alpha = 1 degree, flight 2-37.
181
MO (DEG)
0._ kLl I CI_: [_a ,EL _ -_ IIIIllllllllllllllll IltlllFtr 111111111 IIIIlllllll)[l[ll[ll
,,,, _, , 11 ,, .c IIIJ_ll_[_l_l_LLLlITIllll_*ll_lJllrl_JllllllllJl!
(_F) 0 - S,_'V O-C_G)
o k.0 80 !20 0 _.0 80 12o
i_ SECONDS SECONDS20:W]:02
INo.1io,e_I Imo._io,e_IRB = 7 deg, RC = 9 deg RB -- 7 deg, RC = 5 deg
1.0 ...... %T_LI._qGtN INDE.X_ -_ 1.0- STALLMARGINI,NDE:X
0 .... 0.......................................i.........1
_60. 360CORRECTEDMEIGHTFL.O_ GORRECT1EDV_EIG,HTFLO_
WIR .(PPS) " (PPS) -
o._., i i,_ o_ _ No.s_r._.. I I o.s t _A_O_s,:_,_r__c, No.s_r,_. " "
_AN FAN p_ -F -,--
I :0 JJJ ilJi il_Ji ILIJlJlLI Itllllllf ItJilJlit ILIIJlilf O_ {{{IJ[lll ili_lli:llli[lil L_[I_II[I IliJ_ll il?_li,_
0 WO 80 120 0 40 80 120SECONDS SECONDS
Figure 120. Effectsof sideslipangle on steady-stateinlet characteristicswith SMCV = 20 degreesand alpha = 1 degree, flight= 2-37.
1.82
" ..
Ramp
Wi ng
~Looking aft
Cowl
5 deg
INo. 2 i n1e t IM = 0.84, RB = 7 deg, RCSMCV = 20 deg,~]
Pressure above PTI \in shaded areaPressure below PTJin unshaded area
number onPTI - PT I [1 + ( can tau r )- ~
fUse + if contour in shaded areaUse.- if contour in unshaded area
Alpha =
W1RIDL =
lOT =
IBeta =Beta = 0 degJ
Alpha = 1.4 degW1R = 312 ppsIDL = 0.296lOT = 0.056
!Beta = 1.6 degl
1.3 deg313 pps0.3040.055
AlphaW1RIDLlOT
4.0 deg I1.4 deg311 pps
= 0.4200.076
Dynamic total-pressure contours(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 121. - Effects of S~~l1 wake on no. 2 inlet dynamic total-pressure contours duringsideslip, SMCV = 20 degrees and alpha = 1.4 degrees, flight 2-37.
_0 (DEG)
0
i_ ALTIT_JDE INDICATEDAi_,LE OF SIDESLIP
aLr IT'J_I_ E_TA(v FEET) (DEC,) 0-
,_TT _MCV,.oF) 0 C_E[G)
L_ L _.--_J J-U J- _LLU,--U._U-I
0 20 _0 60 80 100 0 20 _,0 60 80 100
_ 20:5].:[15 SECC_NDS " SECONDS
i.o., 'n'o! ' I"o. 'nlo lRB = 7 deg, RC = 9 deg RB= 7 deg, RC = 5 deg
PT-"_ PTO
0.9-L,L,U..I _.LI L.L .L.L; L_.LL L _:L. L L.L-L . i i J J i i _ i i 1
IDI. I_.
0 IJ,_LL.L L_. _ & _ I 0 I I t I I I I I I I I I I I I I I ] I I _ I I I 11 I ] ] I J I ] I I 1
IDt _ _ J '_ IDT
WlR __ WIR
l_O -_J [_ . I I I I I I I _ I t I _ I 162 I I I I I I I 11 i I I I I I I I I I [ I I I I I I I
o._ ---- FAN Dl SCHARGEMACHNO. SIGtW_L J / 0._t FAN DIS_HI_qGIEI'_C_ NO. S l G/'_/..
1 _'T co_o_) P_" c,u,TCOm_OL)
0 20 '_0 €'0 80 I00 0 20 kO 60 gO tOO
SECOt_S S_CONDS
Figure 122. Effects of sideslip angle on steady-state inlet characteristicswith SMCV-- -8 degrees and alpha = 1 degree, flight 2-37.
[84
Pressure above PTI
I I Wingin shaded area No. 2 inletPressure below PTI
in unshaded area Cowl _ I_ _ Ramp,number on, M= 0 85, RB = 7 deg RC = 5 deg
PTI - PTI [l+ [ contour ) _ ] " '- ^ SMCV = -8 deg Looking Aft
Use + if contour in shaded area
Use - if contour in unshaded area
_J _:_:_:i:_:_:_:i:_:_:_:!:!:i
Alpha = 0.8 deg Alpha = 1.0 deg Alpha = 1.0 deg
WIR = 307 pps WIR = 305 pps WIR = 312 pps
IDL = 0.351 IDL = 0.291 IDL = 0.338
IDT = 0.055 IDT = 0.054 IDT = 0.059
Dynamic total-pressure contours
(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 123. - Effectsof SMCVwake on no. 2 inlet dynamictotal-pressurecontours duringsideslip,SMCV = -8 degrees and alpha = 0.9 degrees,flight 2-37.
ALTITUDE i I(K F[ET) l BETA(Or:(;.)
0 Ill|llll, I''lill,I Illll,l,r llllllll Illlltrl Illlll;I50- 2.5 . .
TE4PfEP.ATURF t i I_[L ON[TROLVANE_POSIT10_NT J(TTO-TTSTO) STRUCT W.._E C
COF) O- (:WEG) I
-50- Iltllllll ZlILfI11_ Zlllllll IIlllltt ll,l(lll IllJllll_ "25 _ _!_ {_.0 80 |20 _+0 80 120
SECONDS SECONDS
_21:).7:52
Imo.' io'e lRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg
PTO PTO
, I III 0 S-- IIIfl IIIIIIiLl I)lllllll Illllllll
..... o0 O"0.2
g I I o,s_'r,o_ I I I 0.2 oIstoRT,o.
0 IIIIIIIIIIIIIIIItllllllllllllllllllltfllllllllllllllllllll{I I iiiiiitlllJflt fll1111111111lllllllllllllllllJllllllfllll_lll j
360 C_tR_CT__IGHTFLO_ 1_60'I , CO.COTEDWEIGHTFLO_ ;
• _.WIR WIR
CPP$) - (PPS) - "
1_0 _llll,llll Illllllll IIIlltll IIIIlllll Ill*tmll _l£1Iltll _60 Illflllll Iltllllll (llllllll {tlffllt tltfl]l] fflllllll0.5
F_ _l_*,_;eA_ NO.s(c.,,_, o.s- , ,FANmsc_ _cHNO.s(c._
O_ IIIIIlll III!llllll Imlllllll I1{111111 IIIJtllll (lllllll 0_ IlZ*lllll iltflrl] i!llllll llfIIIlll flllllll IlmmIItll0 kO 80 120 0 40 80 120
SECONDs SECONDS
Figure 124. -Effects of sideslip angle on steady-state inlet characteristics withSMCV = 0 degrees and alpha = 2.6 degrees, flight 2-37.
186
0- INDI_q'_ A,'_E _ 'ATTACK J
.- HO ALPHA(Dr:6.)
0.5 J-J--LLLJI Illllllll I UJ_.LL I I I _1 II -2 It[
AI_TI TUDE
ALTI'PJDE BL=TA
(K FEET) - (DIEG.) _5-1_ irlllt ,lil,r jiitli ,i!l j0 IlJllllll ililtllll tltlllllf tl[llllll Illllflll
,o_ ,,,,._ "-_----T..... q------r:_-_:-- -1(TTO-TrS_) I- S_UCT_UU.HO0_CONTROLV_E I I
(OF) 0 - (_G) 0 --
._0 f f I I I I I I I I I t I I I I I I ! I / 1111 llZltle:l lillllllll .2_IJ--LL_L_-LL.LL.L_LJ-J_J_J_J-L._JIJ-LI'-ILLJ-._J.
0 20 _,0 60 80 100 0 20 _.0 60 80 100
_, .L._(,ND_< _E_..tlDS
21:12:55
i o., i0,eqRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg
,,,,,.,: ......I _c,,;)L_t v:,,,'LR;
pT_ PTO
,,"_ . li 4_ L_.LJ-:_ j,_ i i_i, ,i ,[,Jl_;,,
IIOT I I _ IOT -r
" 0 iiiiiiiii l 0 I[ ii• 360- .
! . /- -WIR WIR
- - (PPs) CPPS) "
I_0 I I _ I I I I I I I I I I I I l I [ I I , I t i_0 _llllflll _ I I I I I I I I I I I t [ I l J ] ] i I I I I I I I I : I [ I I I I I
0.5_
__ FAN DISCHARGEHACHNO. SIGNAL(AFT CONTROL) (AFT CONTROL)
iII iI II II_; lllI_IllIlllIll!ll lll''_J'l_l
_I_ 01 t I I I I I I I_011 I I I I I I I14_II I I I I I 116(} I I I' I I I I I_(} I I I I I I I I00 _ I I ' I I I I f 120 _O _O SO I_0
SECONDS SECONDS
Figure 125. - Effects of sideslip angle on steady-state inlet characteristics withSMCV = 20 degrees and alpha = 2.6 degrees, flight 2-37.
187
_- Pressure above PTI@o
co in shaded area I I
Pressure below PTI INo. 2 inlet I Wingin unshaded area I I
,number on,PTI - PTI [l +- L contour ) A ] M = 0.85, RB = 7 deg, RC = 5 deg Cowl RampA
r.. • SMCV = 20 deguse + if contour in shaded area Looking aft
Use - if contour in unshaded area
t• .*...,:..,.........:........
_,,_-!...:.!_!
m_-!!
_,"_.i:_:_"/.,,. ,,............ :_.........................._......>.............:..:.:..:.\-_--....-/
I Beta = 0 deg Beta = 1.7 degl I Beta = 3.8 deg I_
Alpha = 2.6 deg Alpha = 2.6 deg Alpha = 2.7 deg
WIR = 317 deg WIR = 317 pps WIR = 318 pps
IDL = 0.391 IDL = O.397 IDL = 0.449
IDT = 0.064 IDT = 0.064 IDT = 0.I04
Dynamic total-pressure contours
(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 126. Effectsof SMCVwake on no. 2 inlet dynamic total-pressurecontoursduringsideslip,SMCV = 20 degreesand alpha = 2.6 degrees,flight 2-37.
i.oi-.---ii...............0 _ UJ./_IItltLJ.J./.J.L /.I.U./_-IJ-L/J I .2_tllllltrlllllllllllllllllll
ALTI'/%IDE BETA .(K FEET) (.DEG.)
i ttllfll iiiii111_ fill11111 i -5 I_Lllllll1_Iti_lllllllltlllll
25.... STRUCT'JRALMOOECONTROLVANE _SITI_ ---
+°- ?1CTTO-TrS_)
ATU -- S/4CV 0- -- --(OF) 0- CDEG) ,
-$0. ]etllrll iil11111 tlllllltlll_tleltl I_1111_t Jllllllsl _25_1LL.L.U.LL_ LJ IJ L.L.LZ.L.L.I-LJ_ JJ.Ai I I I I I _J2I I _, JI I t LL.U.L-U-I-LJ"0 _0 80 120 0 k.O 80 120
_."_ 1$: 57
.g[COI,IDSOk:
I.o.1inletl I"o.2+nlotlRB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg
. 1 II ]jlllll]l 1 I, *,r,,ll,, 11_,,, g._ IIIIII11 ][lllj]ll Illll Illl
GIN INDEX STALL t'ARGIN IND(D(
IDL IDt..
++T+ ' '- (PPS) Wl_
" 160- (PPS)
0..5-- FAN D[_E f_ACHNO. $IGt_4. 0.5- FAN D[SC_GE _4ACHNO. S'|Gt,la,L
(AFT CONTROL) I (AFT cOntroL.) !
51_t_t)_,I_S S_.ONDS
Figure 127. - Effects of sideslip angle on steady-state inlet characteristics withSMCg= 0 degrees and alpha = 5.8 degrees, flight 2-57.
189
-II'{)ICATED An;Lf OF ATTACK V
~-"'" ~~ ___ --.r
-L
~I'{)TCATED AI<;~ OF SIDELIP~
I ,1, ,I
-25
AL_(DEG.)
Sl'CV(DEC)
BETA(DEC.)
_._-~.__ .
SECOI'{)S
ALTln.oE
TEMPERATURE(TTO- TTS TO)
-I--~I
I
- " ~'l
- \0 ~oL.LUl.:..llJ.l..U.J.l..U.J.l..U.~.lJ0Ci.J..L'-ll1-L.LU.:..L..O..L:.s.L1o
J.l..U...LC1.JLU.LL.LU.L:.1lJ
20
~IS:00:52
1.0
110
0.550
ALTlTl.OE(K FEET)
050
~TT
(OF)
120so~O
-~..kIGHTFL~
-
]- ,1J'UliFAN D~SCHAAGE ..lcH I-Cl. Slkw.
(AFT CONTROL)
-
_.1.l.1UJ' u_ .
- -- -:I;~lJ~J~j[Pn-tAX-PfMIN]
PTA\IG
O_ULLLLll-l.11J.JL11.lJLLl-LJLLl-LJ..1.LL:.. l Lt LLl .,
560
0.2 _.
1600.5
lOT
WIR(PPS)
SECOI'{)S SEClNlS
Figure 128. - Effects of sideslip'angleon steady-state inlet characteristics withSMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.
190
SYM TEST p_L_.T/PT 14 ,N..I AI_P_L_ ETA PTO PCNFR-I PCNFR-2 P_NFR-3 PO'IrR-4 I:_tr)UCTION TIME
O 203"/ 6.01 0,87 23059 5.8 0.1 9.'/ 90.04 98.69 82..53 97.06 1/31/'78 2 2 3 18:0:3113 _037 6.02 D.E_ 23015 5.6 1.8 9.6 99.09 99.42 f./_.24 97.35 1/31/78 2 3 2 18:0:51• 203"/ 6.03 0.85 221853 6.2 2.9 9.6 99.13 99.04 6[:1.'/6 9"7.61 1/31/'tB 2. 2 3 18:1:18
.... AIP pressures/PT0 at max IDL - engine l1.040 . i i _ i -
.........................i....................................................i .........................-_..........................-...........................t..........................f ....................................................................................................! ................................................4........................._............................"...........................=:.........................i .......................................................................;::'"": : : . .
........................i.....................................................................................................................................................i........................................................i...........................-- ,.ooo......................"...........:..... ..................................N
o.96o". :i:i.L.....:.:..::::.:.._::]._...i....]..._!.::::&..::ii...........i.:_:i:iLLiiiiiii_ii_if .....L....:._......_.i_ILL...:...:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::iiili i
O. 840 ......................................................._...................._ ..... . .................................................................................................._........................: : s :.. - :- :...........................:............................•............................;.........................._..................................................................................: ...........................,........................... ................_. l " :.
......................................................;...........................'_........................ • .....................................................T...........................'[ .........................T................. T..........................................................._............................_........................._.....................................................i ...........................,............................" .........................." ...........................
o.8oo........................................................i........................_...........].........................................,:......................,;......................_.........................;_............................•: * | . : :" l: : I ....
................................................................................ .:..........................._............................_............................_............................*............................ ...........................-.........................
O.760 ; ! _ ! ! * : :
3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 a 2. 2 2. a 2 2 _ 2 2 2 2 e 2"_l---Probeaaeae21111111J! I000000000099999999998888_'4 3_. I 09 B 76543 e I 098"76543 _ I 098"/ 6543 a I 098 -/6 number
(see figure 104)
AlP pressures/PTO at max IDL - engine 2I.040 _ _ .; i
............................................................................................................•r..........................."...........................f .....................................................-...................................................• ........................ : .........................;..........................,,1...........................,i............................,_........................._ ........................ _,...................... _.....................
,.ooo::..........%'_ ;_:_:_F_.............._f_....._i...........i_:_:J_:_:_:_i-_:_-_-_=_:_:_:-_:_-
iiii iii !-. _'TZ?IIZZf?Z_L..........&LZ22iXl......................j.2ZIZZLZZZ:iZZZZZTZ£
• , . : . : . I .O.800 .........................._............................_............................:............................_.........................'_...................'_......................................................';............ _".............: ; i : : : : "
..........................|............................, ..........................,...........................,...........................,_............................i,............................;,.-_........................, ..........................., ...........................; : ; { : : : : =............................i ...........................;............................;...........................• ..........................._...........................;,..........................._"...........................,i...........................,,7,,..........................
. ! : : [" : " " : • : :
0. 760 : _ ' " _ _ _ ! ' "156666666666666666666666666666666666666663 "_'3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3"3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 386666655555555554444 _ _ 44443333333333_22_=3 4 3 _, I 0 9 8 "/ 6 5 4 3 2. I 0 9 8 'v 6 5 4 3 _ 1.0 9 8 "/ 6 5 4 3 2 I 0 9 8 3' 6
Figure ]29. - IndividualAIP probes, total-pressurerecoveryarray, SMCS vanedeflected20 degrees,Mach 0.85, _-- 6 degrees,flight 2-37.
191
No. inlet
Filtered to 62.5 Hzsampled at 360 sps
No.2 inlet0.10
C0 0.08
+-JL0+-J . 0.06l/l
ICJl"'0
0.lJ't
rc"'0 0.02rc0::: ,
Circumferential distortion
Figure 130.
, ,
Variations in circumferential and radial distortion componentsduring sideslip operation with the SMCS vane deflected+20degrees, Mach 0.85, a= 6 degrees, flight 2-37.
Pres,ure,nsha oabOVearea .... etiI No. 2 inl Wingii
Pressure below PTI
in unshaded area M = 0.86, RB = 7 deg, RC = 5 deg Cowl :_Ramp
,number on, SMCV = 20 degPTI- PTI [l_+ [ contour ) A ]Looking aft
Use + if contour in shaded area
Use - if contour in unshaded area
J
":::i::: _ _" IJ _ _ .1 _ 1 _':"I_ _1_ _, I ,t.::"
:i_iii:_:i_-?i!ii_!??!.".o.I. .j I _. j, i:::i:i:
:iii!!_i:::::::_:_ik;. , .,_ V_:j J::::::::::::::::::::::::
_1 :" ":':".._ :' 1_'""1':':":'1:: : :: :" ":'::::::::: : :" ':':::::."": : : ""
11_ '-" ,-.-....... _ .:.:.:....... t- .:.:: ....... _ 4: •
Alpha = 5.8 de 9 Alpha = 5.6 deg Alpha ; 6.0 deg
WIR = 346 pps WIR = 349 pps WIR = 348 ppsIDL = 0.506 IDL = 0.688 IDL = 0.547
IDT = 0.091 IDT = 0.185 IDT = 0.099
Dynamic total-pressure contours
(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 131. Effectsof SMCVwake on no. 2 inlet dynamictotal-pressurecontoursduringsideslip,SMCV = 20 degreesand alpha = 5 8 degrees flight 2-37° y •
Pressure above PTI Wingin shaded area A
Pressure below PTI T
in unshaded area M = 0.86, RB = 7 deg, RC = 9 deg Ramp _ _- _ Cowl
,number on SMCV = 20 degPTI - PTI [l± t contour ) _ ] Looking aft
Use + if contour in shaded area
Use _ if contour in unshaded area
Beta = 0 deg [Beta = 1.8 deg I iBeta = 3.6 deg[
Alpha = 5.8 deg Alpha = 5.6 deg Alpha = 6.0 deg
WIR = 347 pps WIR = 348 pps WIR = 350 pps
IDL = 0.580 IDL = 0.607 IDL = 0.706
IDT = 0.099 IDT = 0.120 IDT = O.163
Dynamic total-pressure contours(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 132. Effects of SMCVwake on no. 1 inlet dynamic total-pressure contours during
sideslip, SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.
For example,local recoverieson a probe-by-probebasis are illustratedin
figure129. Probeswithin a particularrake are connectedby straightlinesand are arranged in order from hub to tip. This particularplot uses sideslipas the independentparameter. SMCS vane wake ingestionis evidentin theinboardinlet at 2 degreesof sideslipand is becoming evidentin the outboardinlet as sideslipis increasedto 3 degrees. Distortioncharacteristicsresult-
-- Lng from scans that produced the maximtmvalue of stall-marginindex are illu-stratedin figure 130. Combinationsof circumferentialand radial distortioncomponentscomputedat 360 samplesper second are shown for selectedaircraftattitudeswith the SMCS vanes deflected20 degrees. Maximumvalues of stall-
margin index are identifiedand comparedagainst inlet design goals (IDL= 1.0).Correspondingtotal-pressurecontoursare illustratedin figures131 and 132.
Maneuverswith the _CS vane deflected20 degrees (figure133) illustrate
the sensitivityof vortex ingestionto angle of attack. At zero sideslip,vortex ingestionis evidentin the no. 2 inlet betweenangles of attackof7 and 8 degrees,consistentwith the maneuver envelopespresentedearlier.(Seefigure 112.) Similarangle-of-attackexcursionswith the vanes held ina neutral positionare shown in figures134 through136 for comparison.
Steady-stateinlet characteristicsduring sideslipmaneuverswith the SMCSvanes deflected13 degreesare shown in figure137. Total-pressurerecoverydecreasesapproximately5 percentwhen sideslipangle increasesfrom 1.7 to
4.0 degrees. Recovery lossesare accompaniedby significantincreasesindistortionlevels. Total-pressurecontours representingmaximtm dynamicvalues of stall-marginindex as a functionof sideslipangle are shown in fig-ure 138, and a well-developed,low-pressureregion is evidentin the hub.Stall-marginindex increasesfrom 0.5 to 0.75 as sideslipis increasedfrom1.7 to 4.0 degrees.
EngineThrottleTransients. Engine throttletransientswere conductedatselectedconditionsto demonstrateengine stall-margincapabilitybeyond that
requiredby the inlet during SMCS vane wake ingestion. Initialtests were con-ducted during flight 2-19 and 2-33. Qualitativedata and operatingconditionsare summarizedin figuresIII and 112.
During flight 2-38, maximumrate throttletransientsbetween IDLE and-- IhFFEPJ_DIATEpower-leversettingswere conductedon the no. 2 engine during
sideslipmaneuversat differentaltitudeswith the SMCS vanes deflected20 and-8 degrees. Steady-stateresultsat nominalangles of attack between 1 and5 degreeswith sideslipout to 4 degreesare shown in figures139 through141.
195
V'II
RC
in 1et I
F~dIS~~~L(AFT CC,nROL)
I I
I J." '"
, No. 2
O...L.Lu.l.Ju.l.J..l.1..1..l.1..1..l.1..1.Lll.l..Lllll.llLLll
0.2,------,----,------.---, ,-__-,
RB = 7 deg,:.o~
pn.E I 1
PTO
I~L
o•8 ..LI-LJ...LLJ...LU-LLJ...LL..LLJ...LLLLl.LLl.LL.LJ..l..L.
:.0 .,..-----r-----,-----,-
o ..LU..l.1..1.l.l J.l.l..J..LUlllJlllJLl..LLi..LLllLllll.l.ll.l..LllllllllJ..l.l"JJ1.l..l..l..lll..
3&0 F====1'==--==r==JORRECTEO WEIGH; ,~C\, -~1
wlR Ll(PPS)
1&0~ " J0.5
.20SECC1'VS
RECO'/ERY
STALL ",I1'VEX
40 80
I No. in 1et
RB = 7 deg, RC = 9 deg
U",L[uT
LL'.u
L
LL
o\ol1JL.LLJ...LLl..LCLU..l.1..1..LJ...L.lJ
':'LTlTLlJE
JJUl' UlJJJJ " .1.Lllll;LJ
FAN DISC>'AAGE .....CH ,t). SIGNAL
(AFT CONTROL)
=----- I Av
JJJJl
I.
1rmI-- D- ~'"'~~ ,
~ , _111'1!,
50
.J.LTIT1JDE(if. r~ET)
?LA(CE~)
1.0
~PrJ
0.3\.0
IOL
1Q.2
lOT
1HIt)
'~lP
(PP5)
1&0O. ;
FAll ~.Q
200
Figure 133. - Effects of angle-of-attack maneuvers on steady-state inletcharacteristics with SMC vane at 20 degrees, flight 2-42.
196
L CDEO /0
O, _111111!11 iiiiiiiii illlflTll |Tlllltll [llltllll IIit1,111 _ I11 SlI _1!111_
• (K FEET) BETA(DEG)
0 iltllllll Iffllllll IlJtllll! IJllllll I_1111[11111_I[11_11111 Illl[tlt! IlrlrtltllL_0 120 _ _'0 80 120SECONDS SECONDS
RB -- 7 deg, RC -- 9 deg RB = 7 deg, RC -- 5 degL.O ,
EE RECCVERY
p'r_l p'r_PTO PT0
0, 8 Irllllllltlllllll I_111111il I11 r I I IlllllllIl LI II III1 Ill It It i1 0, 8 lilJ_lll rlltlllll Illlfllll
0 Illllll flllll f fill IIIII III II I111111111t II III II II IIIIII1?111 I| Q llJlt IIt IlllIIlll I f III III I III II1111fll[I I11111111111111111 ll|
PTAVG
0 i Ill 0 |llltllll) IIIIllllllllll f IIIIll _1111 flflllljljlflllt II IIIIII/
C_RECT_D WEIGHT FLOW , CORRECTEDWEIGHT FLOW
WIR W1R_PPS) (PPS)
0.5 0._ _ _FAN DISCH/_GE ,-_, NO. STGN_L(_,FTC0NTI_(_,.) (AFT CONTI_OL)
0 Ill|Ill !|lt|f tl|llll! !1|11111 I|11111|1 |1111|1 0 _ I IllJl[I IIIIIIIIJ till|Ill
" IPl..A . PL_
(DEG) (D_G)
0 oIllllrtll IIIItllll Illlltlll Illllll Jllllllll ItllltlllJ 0 _l'_lllll:lllllllllllll_o_.0 0 %20SECONDS sEcOflDS
Figure 134. - Effectsof an angle-of-attackmaneuver on steady-stateinletcharacteristicswith no. 1 engine at intermediate,flight 2-42,
SMCS vane deflectionangle = 0 degrees.
197
IRECOVERY
STALL /'o4ARGIN It-VEx
20
INo.2 inletl
7 deg, RC = 5 deg
I~DICAiE:J .G.l'GL:: 'JF ~TT~CK
..r---..~ ~ lr---
u' .,
2 DISTORTlCJ'.l[PTKAX-PT><INj" PTAVG
1-
, ,
~QRRECTE:> ~ IGHT FLl
5 FAA! OISCHARGE KOoCH NJ. StGN4L(AFT CONTROL)
. !
0' ,
0 F'J1oIEK LEVER fol-GLE
I0
..
1-J
5.,----,------;,..---:-:-::-0------.---,----,{'rCArED""{;L~ ':F ~~1-2Er_SL-I~P_ _+_~__
V
20
II
SECC/-DS
ol60
oliJ.1.l.lll.1.tl.u..J.l..lll.1.Li.J.Jl.UJl..U.Li.J.JLL!.ll..U.J..!..ull.1.UUlll.l.ll.U.l.-'-L
O.
160
O.
lOT
RB =1.0
PTlPT~
0.8
1.0
lDL
eEiA(O~G) k_
WIR(PPS)
PLA(OEG)
GOSECOf,()S
ALTITl.OE
20
~No. inletl
7 deg, RC = 9 degRB~ RECOVERY
~
STALL "1ARGIN [N:lEX
U I~
llll.l.llllu.lllU
IJ .-..-""
·LJ OrSTCRTION
rTKAX-pT><IN]P~iVG
tJI I
C:)RRECTED wE IGHT FLOW
"
F"'N OISCJ-'"Cl.RGE MACH ~. SIGf'.4A.L(AFT CeNTROL)
.~ .f\.A
7.1
1.0
~" ...----r--.....-PQW:-::::ER-L--:~-E"".-t_>-G,...L-E-.--- .....---,
oloO
7
n.2
160
7.5
lOT
ICL
~l~
rpP<j)
1<0 :.r:[un[~_==::-RU'eER
~ u ' ULUllLUlJUlJ..Ll..l..Ll..l..llJ..l..llJ..aJ
-,--c==c:-----r--,------,
'<.T1n;DE'OEfr ... <'"<::::T:
° ll.U llUJ 1 LLlll_l.lllllll.l.JJ.ULLi.J.J..Ll..l..Ll..l.ili.L1.l.l.ll.W
~11:l2:lO
Figure 135. - Effects of an angle-of-attack maneuver on steady-stateinlet characteristics with no. 2 engine at intermediate,
flight 2-42, SMCS vane deflection angle = 0 degrees.
198
\.0 1===+==::[::3""~C~H ~'U'~eE3:'=+===F==l
fAD ').;
ALTlruoE(K FEET)
1+--17:n:2&
L,,~
-< col,> "C" 1---1--,f---+---"-----+-- -, !
I~L
,0
deg
in 1etl
RC = 97 deg,RB
~ J- IRECC"C~Y
lu ,I
"~'l3±JLA~ ~.• o "- "'"
DISTORTION[PT'-'A)(-~""Wll
PT;'VG j
a
IC:JRRE>::::~ rWETSI-" .:'L1W
a ~.
\ FAN 'JISCI"¢.RGE "'ACH f.IJ. SIGNAL(AFT CCt,TROL)
!~
'J "
POtiER L::~E~ ..:.tGLE
If\
",.
36
~ 6a.
O.B\.0
200
a0.2
\.0
IDL
lOT
~PTa
WIR~PPS)
FAN ff
SECCtJ)S
Figure 136. - Effects of an angle-of-attack maneuver on steady-stateinlet characteristics with no. 2 engine at intermediate,
flight 2-42, SMCS vane deflection angle = 0 degrees.
199
Figure 137. Effectsof sideslipangle on steady-stateinlet characteristicswith SMCV = 13 degreesand alpha = 3.3 degrees, flight 2-38.
200
Pressure above PT I I I _Wing
in shaded area No. 2 inlet
Pressure below PTI
in unshaded area M = 0.85, RB = 7 deg, RC = 5 deg Cowl -_4_>- Ramp
,number on SMCV = 13 degPTI - PTI [I+_t contour ) A ] Looking aft
i
Use + if contour in s'haded area
Use - if contour in unshaded area
I lBeta = 0 deg Beta = 1.7 deg Beta = 4.0 deg
Alpha = 3.4 deg Alpha = 3.2 deg Alpha = 3.2 deg
WIR = 348 pps WIR = 347 pps WIR = 347 pps
IDL = 0.514 IDL = 0.504 IDL = 0.750
IDT = 0.090 IDT = 0.099 IDT = 0.163
Dynamic total-pressure contours i
(max IDL scans, filtered to 62.5 Hz, sampled at 360 sps)
Figure 138. -Effects of SMCVwake on no. 2 inlet dynamictotal-pressurecontours during•o'_ sides-lip, _ICv'= 13 degrees and alpha = 3.3 degrees, flight 2-38.
&0
SECCNlS
20
-aDlCATfD At-&L~ OF ATTACK
"- ~ V'-- '- ~-v-v
_1-ll.LLJ..ll.I.JL.U.1.LU.LLJ..WJCLl..l.LU.LLJ..ULWJ.LU.LUULWJl.U..1.LU.ll.LJ
ALPHA(OEG.)
BETA(OEG.)
&0SECCN>S
o 20
~20:00;4b
[1MAC.H N..f"8ER
-- -_.
-ALTlruoE
-
LLLU
0.1\0
1.0
MO
:"I.TlflKJf(r FEET)
RB
in 1etl7 deg, RC = 9 deg RB
INo. 2 in 1e tl7 deg, RC 5 deg
&0
STALL MARGIN I/{)f)(
20
WJ"(PPS)
IOL
PTIPTe
lOT
PLA(DEG.)
&04020
:rRECOVERY r
'\l/\-vv -v
STALL MARG IN IIllEX
) 1"4 V -"'y- rv-
- U
r--
JIV'I n..,.~ '\OIJ:I~-~ [PTl<AX-PTMIN]
PTVG
-I--- or
~RECTED 11GKT FLJ-
,,"~LJ,~_ JJUJ1..l.l-l-
FAN DISCt-'ARGE r-oACH tD. SIGNAL(AFT CONT"OL)
- Iv-.
--.---
PCWER LEVER AAGLE
.r--t-------.
0o
1.0
o0.2
o1&0
0.81.0
IOL
lOT
~PTO
WI'(PPSI
1&00.1
FAN ¥
0200
PLA(;lEG. )
SE(CIllS SECCN>S
Fib~re 139. - Steady-state lnlet characteristics during no. 2 enginetransients, sMcv ~ 20 degrees, alpha = 5.5 degrees,and beta = +4 degrees, flight 2-38.
202
S-r----,-----r----,-----,----,
", ]~--'~t[l~::~>]""""g,~','~,blJ
a 20 40 60 80 100
I SECOI'()S
f-----21:SI:S8
INo. lin I e tl
.oLPHA(DEG.)
BETA(DEG.)
8!>Il!CATED Al'GLE OF ATIACK
a-2- ."
I1>IliCATED Al'GLE OF SIDESLIP
I20 80
SECOI'()S
INo.2 inletl
lOa
o ' . , "-Ll.LU..LU..LU..LU..u..L.u..L.u..L.w
200_
l:- """Ell LEYOl AN::LE
- r-r-'-Li~- - --\.-, r'-
_ ll1.-t ' 'I
'~IR
(PPS)
FAN ~p
PeA(DEG. )
20
SEem,DS
60 80 lOa
IDT
WIR(PPS)
F/>I< -¥
PLA(DEG.)
SECCKJ5
Fi~lre 140. - Steady-state inlet characteristics during no. 2 enginetransients, SMCV = 20 degrees, alpha = 3 degrees, andbeta = +4 degrees, flight 2-38.
203
240160
I--l rulm~DiATEIDLE
H""tE.,;: Lt.VF.R MoGLE
"\. .J rvREJVERY
.. ~.
-
-5. ~. 160 240
SECDt<lS
20 deg -8 deg •INo. 2 in 1etl
RB 7 deg, RC =-5 deg1.0
PLA(oEG.)
240
BETA(OEG.) 0
240
ISMCV
160
SECQr-()S
SEcoms80
FAN DISCtAAGE r-tACH N>. SIGNAL(pET CONTROL)
- I.
--t Pl"MER LEVER A1'C.LE
-lrJI~~\LJlfL-.-.-lJU
_Wlll.llll ,
AL_(oEG.)
1.0- .._- u_ r----.---~--..._--__,___--__,
o'00
o 0
o
l---21: 40: 48
'No. 1 in 1etlRB 7 deg, RC = 9 deg
1.0
~ PTIPTO ?TO
0.'1.0
101. 101.
00.2
lOT lOT
PLACCEG.)
MO
WI"(PPS)
AL TIT\X:lE(<FEET)
Figure 141. Steady-state inlet characteristics during no. 2 enginetransients, ~1CV = 20 degrees and -8 degrees,alpha = 1 degree with positive sideslip, flight 2-38.
204
SU_V_RY
B-I experiencefrom air-inductionsystem flight tests with vorticesgeneratedby the SMCS vanes is summarizedwith the followingobservations:
_ (i) Agreementbetweenflight-testand wind-tunneltest resultswasgenerallygood. Small differenceswere noted in maneuvers (combinationsofangles of attack and sideslip)that resultedin wake ingestion.
(2) Inlet total-pressurerecoveryand distortioncharacteristicscan beadverselyaffectedby wake ingestion,and considerationof these factorsplaysan importantrole in establishinggood inlet/enginecompatibility.
(3) Tests were conductedwith the vanes deflectedand held at a staticconditionand also with the vanes driven plus and minus full-scaledeflectionat a frequencyof 1 Hz. Tests with oscillatingvanes verified SMCS vanewake/vortexingestionin flight. Tests with staticvane deflectionsproducedresuitsconsistentwith oscillatingvanes and were less objectionabletothe crew.
(4) SMCS operationhas not resultedin any flightrestrictions. Dynamicvalues of stall-marginindex remainedwell within allocationsfor all conditionsinvestigated,and there were no flight incidentsrelated to the conductof theSMCS tests.
(5) Effectson total-pressurerecoveryand engine-facedistortionweregenerallyrestrictedto windward sideslipoperationcombinedwith leading-edge-upvane deflections. The 13-degreevane deflectionangle produced effectssimilarto the 20-degreedeflectionangle on inlet performance.
(6) Sideslipangles resultingin wake ingestionare smallerduring opera-tion at higher anglesof attack due to the increasedoutwash from the fuselageforebody. Also, the range of sideslipangles resultingin wake ingestiondecreasesas angle of attack increases. No wake ingestionwas encounteredabove 7.5 degreesangle of attack.
S_Y OF SMCS FLIGHT TEST.RESULTS
" The objectivesof the SMCS flight-testprogramwere fourfold: (i) obtainspecificdynamicresponse data to validateanalyticalmodels of the aircraftand controlsystems, (2) determinedetail SMCS performancecharacteristics,(3) determineimpactof SMCS on handlingqualities,and (4) determineimpactof SMCS on operationalcapabilitiesof the aircraft.
205
l'ouraircraftwere built under the B-I contract;A/C-I and A/C-2 wereused for detailedengineeringand developmenttests while A/C-3 and A/C-4were used for operationaltests. As of 16 March 1979, B-I SMCS-relatedtesttime was as Follows:
A/C-I 13.6 hr
A/C-2 3.3
A/C-3 116.5
A/C-4 0.3
Total 133.7 hr
A considerableamount of the flight-testdata has alreadybeen discussedin associationwith the topicsof previousparagraphsof this report. It istileintentof this sectionto present flight-testdata not touchedupon by thespecifictopics coveredearlier.
_S PERFORMANCE IN TIME-HISTORY-DATA FORMAT
Accelerationtime historiesare perhaps the most dramaticmeans of demon-stratingthe ride qualityproblem solvedby the SMCS. In the vertical axis,the syn_etricfirst-fuselagemode at approximately3 Hz is the big motionproducerat the nose of the aircraft due to turbulence. In the lateralaxis,tileantisymmetric(lateral)first-fuselagemode at approximately5 Hz producesthe largestresponseto turbulence. Figure 142 illustratesthese motionswith typicaltime-historyplots of verticaland lateralaccelerationat thenose of tileaircraftat the SMCS vane location. The aircraftwas flyingthrough low-altitudeturbulenceat M = 0.75with the SMCS off at the timethat these recordswere taken.
Figure143 shows data similarto that of figure142; only during thisflighttest, the SMCS was operatedto determineits effectiveness. The datasl_cn were recordedon a-flightat M = 0.70 where the B-I was flying ataltitudesof 305 to 610 meters (i,000to 2,000 feet) above the terrainin thelocal EdwardsAir Force Base area. Considerablelight to moderateturbulencewas presentnearly continuously. The SMCS was turned on with both the vertical -
and lateralcockpitgains setat 1.5. To demonstratecomparativeaircraftper-formancewith and without SMCS operating,severaltime periodswith the SMCSon and off were recorded. Figure 143 is typicalof these data. The pilotstationverticalaccelerometerwas not operativeduring this flight,so thereadingof the vertical accelerometeron the radome is shown.
2O0
A2008
Vertical acce] 3 Hz first fuselage symmetric
at FS 572 (225) (vertical) mode
+0.129 g
•1.0 g---0.129 g
_k
+0. 132 g
0 g---0.132 g
-
A20I ILateral acce] 1 5 Hz first fuselage antisymmetric (lateral) mode i
i sec
SCAS on, SMCS off
M = 0.75, alt = 152.4 m (500 ft)
Figure 142. - Typicaldynamicresponse near crew station dueo
to turbulenceduring low-altitude,high-speedflight.
F,;COo
SMCS off SMCS on
Radome uE "- _ Radome E ._
vertical ._ vertical _ _ [velocity _ velocity "_._A2027 r-,_
Fs56(22)_ _
Pilot- us Pilot- u m
stationlateral _ station m c_.• lateral °i_
velocity _JL _ +_A2011
FS 747 (296) Right ° I _ ^ . . ^ Avane F-,
dell ,_i
Left °
vane
defl
Figure 143. SMCS performancein turbulenceM = 0.70, _t = 305 m (I,000ft) AGLA= 65°•
!
Verticaland lateralmotion at the front end of the aircraftwith the
SMCS off are shown by the first two time historiesof figure 143. As indi-cated, the primarymotion in the vertical axis was the first fuselagebendingmode at approximately3 Hz. The lateralmotion was composedof whole-vehiclemotion near 1 Hz and the first-fuselageside-bendingmode motion of approxi-
_ nmtely 5 Ilzsuperimposed. When the SMCS was operated,as shown in the nexttwo plots, considerableattenuationof the 3 Hz motion was achieved. Verylittlemotion of the aircraftat lower frequenciesappearsto be present. Theeffectof the SMCS on the lateralaxis motion was not as dramaticas on the
verticalmotion, but the 5 Hz motion was partiallysuppressed. The whole-vehicle lateralmotion was not attenuated. It is to be recalled,however,that the SMCS is designedto attenuatestructuralmode responsewithoutadverselyaffectingw|mle-vehiclemotion (handlingqualities). The last twoplots in figure143 show the SMCS right and left vane motion during the timethat the SMCS was operating. As shown,the maximumvane deflectionsseldomexceeded 16 degrees,whereas, +20 degreeswere available. Both the 3 Hzvertical and 5 Hz lateralstructuralmotion can be seen to drive the vane
deflections;the largestcomponentis due to the verticalmotion.
Anothermeasure of the SMCS effectivenessis the amount of structural
dampingthe system is able to provideto the key fuselageresponsemodes. Theupper left-handplot in figure 144 shows that it was possible to excite thes_unetricfirst-fuselagebendingmode with a sharp horizontal-tailinput pulse.;:tomthe time historyof the vertical load factor at the SMCS vane location(afterthe horizontal-tailpulse was removed),it was possible to extractthestructuralmode dampingratio, _. Figure 144 shows a plot of _ obtained inthismanner versus SMCS verticalgain setting _as set in the cockpit). A
typicnlnominalgain settingof 1.5 is indicated. The upper right_handplotshows the time-historyresponseof the normal load factor with the verticalgain at this setting. Figure 145 demonstratesthe vertical SMCS performanceover a wide range of vehicleweights.
Attemptswere made to excite the antisymmetricside-bendingmodes withsharp lower rudderpulses and to extract structuralmode dampingratios.
- This did not prove to be a successfultechnique. A differentexcitationtechniquewas used. As has been mentionedin an earliersection,A/C-Iand A/C-2 of the test B-I aircrafthave syst_nsinstalledallowingthe SMCSvanes to be oscillatedat variousamplitudes (A) across a range of frequenciesup to i0 Hz. This capabilitywas used in an attemptto extract structuralmode dampingfor lateral side-bendingmodes. The techniquewas to selectaresonantfrequencyand amplitudesufficientlylarge so asto providetransientdata when the forcingmotion was cut off. The data of figure 146 shows thatit was possibleto excite the 5 Hz first-fuselagelateralbending mode rathercleanly. .Exceptfor SMCS off response,the decay responsecould not be usedto extractthe structuralmode-dampingratio because of high-frequencymode
209
SMCS off SMCS on, vert gain 1.5
! I t I, I ! _ I
0 ] 2 3 0 l 2 3
nz Time- sec nz /'_ Time- sec
FS 572 FS 572
(225) (225)
n
z nzFS 2649 .._ FS 2649
(IO43) (I043)
LH 6 cv J_
RH _ cv /__
j-0.15
Structural I
mode O.lOtime history
damping 0.05ratio plots
0 , , , , I . • _ , • I0 l 2
SMCS vert gain setting
A = 650 , M : 0.85, alt = 914 m (3000 ft)
Figure 144. - First fuselagesymmetricstructuralmode dampingfrom horizontaltail pulse excitations.
oM = 0.85, alt = 18i9 m(6000 ft), wt = 131 544 kg (290 000 Ib)
, 0 M = 0.85, alt = 914 m (3000 ft), wt = I16,575 kg (257 000 Ib)
O M = 0.85, alt = 914 m (3000 ft), wt = 123 969 kg (273 300 Ib)
OM = 0.85, alt = 914 m (3000 ft), wt = 144 698 kg (319 000 Ib)
M = 0.85, alt = 152 m (500 ft), wt = 127 008 kg (280 000 Ib)
A= 65°
Structural mode frequency,3 Hz
0.20
I
IO. 15 f
structural Imode 0.l0 I
damping Iratio il
Nominal
0.05 Settingg
II
" 0 I I I I I0 0.5 1.0 1.5 2.0 2.5
SMCS vertical gain setting
Figure 145. Effect of SMCS verticalgain settingon first fuselagesymmetricstructuralmode dampingat variousaircraftweights.
211
Lateral acceleration Exciter --
at SMCS vane FS 572 (225) off SCAS off, SMCS offI
'±0.J8 II
A=150 I
f=5 Hz I SCAS on SMCS off -Ii
+-0.20
A=150 i
I SCAS on, SMCSiI Lat gain 1.0
A=I50 II SCAS on, SMCSII Lat gain 1.5
+0.15 gl I
A=200 II
I SCA_Son, SM_C-SI! Lat gain 2.75
+-O. 18 g_I
A=300 I _ l sec-_
A= 65°, M = 0.85, Alt. = 762 m (2500 ft)
Figure 146. First fuselageantisymnetricstructuralmode dampingfrom forced SMCS vane oscillations.
212
contamination. The data do indicate that the SNCSis effective in eliminatingstructural response motions at the lateral system gains tested.
SMCS PERFORMANCEIN PSD-DATAFORMAT
in additionto time-historydata, anotherconventionalway of lookingat
ride qualityperformanceand the effectof the SMCS on performanceis in theform of PSD plots of load factorsat the pilot station. Figures147, 148,and 149 show typicaldata of this type for the vertical and lateral loadfactorswith SMCS off and on. Becauseof other test requirements,none of
the B-I flight-testaircrafthad a gust boom installed. Approximateverticaland lateralgust intensitieswere estimatedusing nose-boomangles of attackand sideslipangles in order to providenormalizingfactors for the PSD data.Gust intensitiesfor the data shown were estimatedat 1.22 to 1.52 meters per
second RMS (4 to 5 feet per secondRMS).
l:igure147 shows that the approximately3 Hz first-fuselageverticalbendingmode, previouslyshown in the time-historyplots of figures 142 and143, ks the main contributorto the verticalmotion at the pilot station. Asin figure143, the data of figure147 demonstratethat the SMCS is veryeffectivein reducing the pilot vertical load factor.
PSD plots of the pilot stationlateralload factorare shown in figures148and 149 for the SMCS off and on conditions. Insteadof a consistentsingle-
peak responseas in the vertical case, the lateralresponse exhibitstwodifferenttypes of responses,dependingon fuel loading. Figure 148 illustratesa single-peakresponse as seen in the time-historyplot of figure 142. Fig-gure 149 shows the other common responsewith two peaks between 4.5 to 6 Hz.In this latter instance,the time,historydata appearmore random than thelateralaccelerationtrace of figure 142. These data were taken with the SMCSforwardsensor package in its originallocation. The SMCS is seen to signifi-cantly reducethe main responsepeaks but tends to excite some of the higherfrequencymodes in the immediatevicinity. It was this couplingthat led to
. the sensor relocation study discussed in detail earlier.
-- SMCS AND HANDLINGQUALITIES
One of the design goals for the SMCS was not to interferesignificantlywith basic handlingqualities. The impactof SMCS operationon the B-I handlingqualitieswas determinedfrom horizontal-tailand rudder-doublettransientresponses. The resultsof these tests are shown in figure 150. The wingswere at the 65-degreesweep position,Mach 0.85, and altitude1524 meters(5000 feet). During these tests, the CG locationwas varied.
213
SCAS only
SCAS + SMCS (cockpit gain set 1.5)
--4
0.00025
0.0025
0.00020
0.0020 _nz
0.00015
g2 0.0015' g2
(m-_sec)2 (ftlsec) 2tad/see
rad/sec 0.000100.0010
0.0005 0.00005
0 - 0 _0 lO 20 30 40 50 60
Frequency - rad/sec
I, I I I I I I I I I I
0 l 2 3 4 5 6 7 8 9 lO
Frequency - Hz
Wt = 125 194 kg (276 O00 Ib), CG at .35 _w
= 65°, M = 0.85, alt _152 m (500 ft)AGL
Figure 147. - Vertical SMCS performance in turbulence as shown by power spectraldensity of vertical load factor at pilot station, FS 747 (294).
214
SCAS only.... SCAS + SMCS (cockpit gain set 1.5)
0.40 xlO--3
"- --L
0.35 xlO
0.35-
0.30_ny
o.30- -2Jg2
(m/sec)2 O.25rad/sec g2
0.25 -(ft/sec)2rad/sec
0.20
0.20 -
0.150.15-
0.100. I0 -
i¢'_\
I \€'\
0.050.05 - /\\ ¢/ _., , \\
O- 0 _ -_ _-. -_-0 10 20 30 40 50 60
f
Frequency _ rad/secI I I I I I I I I I I
-. 0 1 2 3 4 5 6 7 8 9 10
Frequency-, Hz
Wt = 125 194 kg (276 000 Ib), CG at .35 _-w
A= 65°, M = 0.85,alt_ 152 m (500 ft)AGI
Figure 148. - Lateral SMCS performance in turbulence as shown by power spectraldensity of lateral load factor at pilot station,
FS 747 (294), single-peak response.
215
SCAS only
SCAS + SMCS (cockpit gain set 1.5)
0.12
0.12 m xl0 -4
xl0 -3
0.I0
0. I0
0.080.08
g2
(m/sec)2 g2 !I 13
rad/sec (ft/sec)2 II
rad/sec I1t_
0.04 0.04 \/ _
/ 1
O. 02 O. 02 I I r_\ / "x
../ / I,","x
0 _ OJ0 l0 20 30 40 50 60
Frequency- rad/sec
i I I I I 1 I I I I I0 1 2 3 4 5 6 7 8 9 I0
Frequency - Hz
Wt = 128 369 kg (283 000 Ib), C.G. at 0.46 Ew
A= 65° ' M= 0.70, alt _305 m (1000 ft)AGL
Figure 149. - Lateral SMCS performance in turbulence as shown by power spectraldensity of lateral load factor at pilot station,
S 747 (294), double-peak response.
216
25" °Dutch roll mode 2.0
frequency - 1.5rad/sec
l.O
.5
0 I I I I I0.2 0.3 0.4 0.5 0.6
C.G - MAC
5-Q Q
Short-period mode --_--B
frequency - 3
rad/sec
21 _SCAS on, SMCS off--O--SCAS on, SMCS on
!
0 I I I I I0.2 0.3 0.4 0.5 0.6
C.G.-MAC
Wt = I17 482 kg(259 0001bs) to 128 822 kg(284 O001bs)
A = 65 °, M = 0.85, alt = 1524 m (5000 ft)
Figure 150. - SMCS impacton short-periodand dutch roll frequencies.
217
In the longitudinalmode, short-periodmode frequencyand dampingratiowere extractedfrom the transientfollowingthe pitch doublets. A slightreductionin the short-periodmode frequencywas observedwhen operatingwiththe SMCS on (figure150) for all CG positions. No significantchange in theshort-periodmode damping ratio was observed.
In the lateral-directionalmode, the Dutch-rollmode frequencyand damping --ratio were extractedfrom the transientfollowingrudder doublets. Resultsobtainedfor the lateral-directionalhandlingqualitieswere similarto thosefor the longitudinalhandlingqualities. The Dutch-rollfrequencyis slightlyreducedby the SMCS at all CG _ositions. (See figure150.) Again, no signifi-cant effectof SMCS on Dutch-rolldampingcould be detected.
In order to determinewhether SMCS would interferewith aircraftmaneuver-_ng,a roller-coastermaneuverwas executed with the SMCS off and on. Evalua-
tions of recordeddata of the SMCS vane deflectionsshowedno significantmotion;maxin_mvane deflectionsrecordedwere less than 2 degrees. This wasdeterminedto have a negligibleimpacton the maneuveringof the aircraft.Subsequentto the previouslydescribedtests, pilots have reporteda slightincreasein stick force required in terrain followingwith the SMCS on overthat with the SMCS off.
SMCS HICH-GAINTESTS
B-I A/C-I and A/C-2 were used in the high-gaintests to demonstratesyst_gain marginsover the expectednominal gains. It had been initiallyplannedtodo all of the testingof the relocatedforwardSMCS sensor packageon A/C-Ionly, includingthe high-gaintests under discussion. However, after only afew of the tests associatedwith the relocatedforwardSMCS sensorpackagehad been completed,A/C-I went into layup for modifications. In order tocontinuewith testing,the forwardSMCS sensor packagewas relocatedonA/C-2.
A summaryof the high-gaintests on A/C-I and A/C-2 is presentedintable X. At nominalgains, there appearsto be no differencein the SMCS
performanceon A/C,1 or A/C-2. At high gains, however,A/C-2 wan limitedbya 35 Hz limit cycle which was not evident in similarA/C-I data. These data _-show that A/C-I and A/C-2 have nearly a factor of two gain marginsoverexpectednominalsettingsfor the ride qualitydesign point of Mach 0.85 at
low altitudes. It is at Mach 0.55 at low altitudeson A/C-2 in a lightweightconfigurationthat the 35 Hz limit cycle preventsobtainingthe expectednominalgains. From the lightweightconfigurationdata of flight1-41 obtainedat Mach 0.85 at high altitudewhere the dynamic pressure schedulegain in the
218
• o i I
TABLE X. - SMCS HIGH-GAINTEST_ SUMMARYi
WingHeavy weight Light weight
sweep M = .85,low alt M = .85 low alt(degrees)
65 Vertical Lateral Flt Vertical Lateral Flt
gain gain no. gain gain no.
1.9 2.2 1.9 2.2
expected expected expected expectednominal nominal nominal nominal
4.8 3.9 2-23 4.5 6.0 1-48
35 Hz limit cycle no 35 Hz3.0 6.0 a2-25
35 Hz evident
Heavy weight Light weightM = .55, low alt M = .55, low alt
55 Vertical Lateral Flt Vertical Lateral Fltgain gain no. gain gain no.
3.0 3.7 3.0 3.7
expected expected expected expectednominal nominal nominal nominal
4.5 4.7 2-23 3.0 3.1 2-23
35 Hz limit cycle 35 Hz limit cycle
S.0 5.5 bl-50no 35 Hz
ia,Repeat of 1-48 runsbRepeatof 2-23 runs
SMCS is about the same as at the low-altitudemach @.55 flight condition,itwas expectedthat a verticalgain of six would be obtained. However, as thetable shows,this value was not obtained.
These inconsistenciesin A/C-I and A/C-2 data led to repeatingwith A/C-2a set of previouslyrun A/C-I tests and vice-versa. These tests were completedon flights 2-25and 1-50. Analyses of these data indicatethat the 35 Hzgain limitationsare unique to A/C-2. Ground tests were conductedon bothaircraftin an attempt to identifythe causes of the 35 Hz limitationsofA/C-2; these tests were unable to isolatethe causes of the 35 Hz.
CREW EVALUATIONSOF SMCS EFFECTS
The bulk of the SMCS operationalsuitabilitytests were conductedon B-I
A/C-3. Ten B-I flight crewmenparticipatedin the tests: four pilots (PLT),three flight-testengineers (FTE),and three OffensiveSystemOperators (OSO).The crews flew regularlyscheduledmanual terrain-following(MFF)and automaticterrain-following(ATF)flight-testmissions with the SMCS on and off. Crew-
memberswere instructedto maintain awarenessof comfortand personalper-formanceduring the various TF missions. Each crewmemberwas instructedtocomplete a test questionnaireto documenthis evaluationof the B-I ride andeffectsof this ride on his performance.
The curves in figure 151 through158 summarizethe subjectiveresponsesto questionson overall ride quality and the effect of turbulenceon: flightpath/nonflightpath controltasks; readabilityof instrumentsand displays;reaching/usingcontrols;crew fatigue,motion sickness,and physicaldiscom-fort. These data are in the form of compositeresponsesfrom all crewmembers.
Overall Ride Quality
Subjectswere asked to rate ride qualityduring TF flight as a functionof SMCS on and off for the followingconditionsof turbulence: smoothairand light,moderate,and heavy turbulence. The combinedratings for all Lsubjectsare shown in figure 151. The data reveal the following:
(i) Ride quality ratingsdecrease (ride qualityworsens) as turbulenceincreases.
(2) In smooth air, use of the SMCS results in little improvementinride quality.
(3) Use of the SMCS improvesride quality in light,moderate,andheavy turbulence.
220
N=7Poor- \ _ SMCS off
a"a,
c_ N = lO s N = lO
'_ Fair- SMCS on 4 PLTCD
3 FTE
_. 1,/ 30SO4J _ p
m Good- ,,$II N = 7 N = 7• _ •_S S
• cr ." 4 PLTN= I0
-o 2 FTE
Excellent -- l OSO
N = number of test
I I I I participantsSmooth Light Moderate Heavy
Turbulence rating
Figure 151. Ride quality ratings for varying degrees of turbulence.
Turbulence effect ratingN = 4 PLT
_orkload Performance
effect effect
Unable to perform task-
Large-- i _
Max effort Major--_. SMCS off
Moderate-- _
Moderate-- __ SMCS on
• Incr effort Minor--None --
- None [ None--t
I I ILight Moderate Heavy
Turbulence rating
Figure 152. - Effectsof turbulenceon flightpath controltasks.
221
Turbulence effect ratingN = lO
Workload Performance 4 PLT
effect effect 3 FTE .
Unable to perform task- 30SO
I Large - N = 7
Max effort Major-- N = 7 4 PLT
2 FTEModerate- N = I0 _-" SMCS off 10SO
{Moderate--_ _" N_ SMCS on_s.. _ '
Incr effort Minor-- f
None- = 7
NoneI None - N = lOI.
I I ILight Moderate Heavy
Turbulence rating
Figure 153. Effects of turbulence on tasks other than flight path control.
Figure 154. Effectsof turbulenceon readabilityof instrumentsand displays.
222
N=IOExtreme --
4 PLT
3 FTEc N=7
-*_ Moderate- _,,_ SMCSoff 3 050
11 N = 7
N = lO 11.1_ SMCS on
_./// ..__/_ _. 4 PLT\._u Slight-- X_ /////- \ 2 FTE
_ .//" % ..= 7 I OSO
None -- N = 10
I I I ISmooth Light Moderate Heavy
Turbulence rating
Figure 155. - Effects of turbulence on reaching/using controls.
Extreme-- N = 10N=7
c \ 4 PLT
"_ -_ ...- SMCS offm _-" 3 FTE
Moderate-- ._ SMCS on 30SO4-J
u N= I0 N=7
_- 4 PLTSIight --
. o N = 7 2 FTE10SO
"_ N= 10None --
I I I ISmooth Sight Moderate Heavy
Turbulence rating
Figure 156. - Effects of turbulence on crew fatigue.
223
c_ N= I0c Extreme-
SMCS off 4 PLTrO
SMCS on 3 FTE4-'
PIoderate- 30SOtl..,,
_- N=7iz1
N = lO 4 PLT
Slight- 2 FTEu 10SOt/3
r-
.o None -4J
0
I I I lSmooth Light Moderate Heavy
Turbulence rating
Figure 157. Effects of turbulence on tendency for motion sickness.
_- N = lO._Extreme --
fDL SMCS off 4 PLT
SMCS on 3 FTEU
_- Moderate- 30SO
N=74--1L.
o 4 PLT
OE Slight -- _ _ _k_1, 2 FTE'_ "f _'_ I oso ":g
None- N = 7 "U
U3 -->..¢-" I I I I
Smooth Light Moderate Heavy
Turbulence rating
Figure t58. El;fectsof turbulence on physical discomfort.
224
One subjectadded the followingcommentto his ratings: "In smoothair,SMCS is not needed. Standardof comparisonis the F-Ill which is excellent."(OSOA)
Flight-PathControlTasks
Pilotswere asked to rate the effectsof turbulenceon flight-pathcontroltasks during TF with SMCS off and on. Ratingswere to be made with referenceto a turbul_ence-effectrating scale. The combinedratingsfor all subjectsare shown in figure152. The data reveal the following:
(i) The effortto performflight-pathcontroltasks and the negativeeffect on task performanceincreaseas turbulenceincreases.
(2) Greatereffort is required,and the negative effect on subjectperformanceis greaterwith SMCS off than with SMCS on; i.e.,workload isless When SMCS is being used.
Non-Flight-PathControlTasks
All subjectswere asked to rate the effectsof turbulenceon non-flight-path controltasks during TF with SMCS off and on. Ratingswere made to aturbulence-effectrating scale. The combinedratingsare shown in figure153.The data reveal the following:
(I) The effortto performnon-flight-pathcontroltasks and the negativeeffecton task performanceincreaseas turbulence increases.
(2) Greatereffort is required,and the negative effecton subjectperformanceis greaterwith SMCS off than with SMCS on; i.e., workload is lesswhen SMCS is being used.
One subjectadded the followingcogent, "Operationof equipmentdifficult-particularlyCITS." (FFE C)
Readabilityof Instrumentsand Displays° -
All subjectswere asked to rate the degreeof difficultythey experiencedin reading instrunentsand displaysduring TF flight as a functionof SMCS offand on for four conditionsof turbulence: smoothair and light,moderate,and
225
heavy turbulence. The combined ratingsfor all subjectsare shown infigure154. The data reveal the following:
(1) The difficultyto read instrumentsand displays increasesasturbulenceincreases.
(2) Difficultyratingsare higher (readabilityless difficult)with SMCS ""on than with SMCS off.
One subjectadded the followingcon_nent:"'E' scope always extremelydifficultto read because of location - turbulenceno factor."
Reaching/UsingControls
All subjectswere asked to rate the difficultyin reachingcontrolsorin performingcontrolactionsduringTF flight as a functionof SMCS off andon for four conditionsof turbulence: smooth air and light,moderate, andheavy turbulence. The combinedratingsfor all subjectsare shown in fig-ure 155. The data reveal the following:
(i) The difficultyto reach and use controls increasesas turbulenceincreases.
(2) As turbulenceincreases,it is easier to reach and use controlswithSMCS on than with SMCS off.
Five of the subjectsadded a conlnent,as follows:
(i) "In manual TF, pitch stick force too heavy." (PLT C)
(2) "Ratingis based primarilyon use of central integratedtestsystem (CITS)."(FTEA)
(3) "CITS controlssomewhatdifficultin turbulence." (FTEB)
(4) "Operationof CITS is difficultat best. SMCS helps." (FTE C)
(5) "Some of the ratingsare due to the locationsof the controls."(OSO C)
226
Crew Fatigue
All subjectswere asked to rate the degree to which TF flight introducedany specialtendenciestoward fatigue. Ratingswere requestedfor the SMCSoff and on modes for four conditionsof turbulence: smoothair, and light,
_ moderate,and heavy turbulence. The combinedratings for all subjectsareshown in figure 156. The data reveal the following:
" (I) Fatigue effectsincreaseas turbulenceincreases.
(2) There is little differencein fatigue-effectratingswith SMCS offversus SMCS on except for the moderate turbulencecondition. For moderateturbulence,the fatigue-effectsrating for SMCS on is higher (less fatigueeffect)than for SMCS off.
Three of the subjectsadded a con_nent,as follows:
(i) "Ratingsmade for AUFO TF." (PLTA)
(2) 'Turbulenceis not the primary factor in producingfatigue - TFitselfproducesa high level of fatigue." (FTEB)
(3) "Turbulencedefinitelyincreasesfatigue." (FTEC)
Motion Sickness
All subjectswere asked to rate the degree to which TF flight introducedany specialtendency for motion sickness. Ratingswere requestedfor theconditionsof SMCS off and on for four levelsof turbulence: smoothair and
light, moderate,and heavy turbulence. The combinedratings for all cre_nenare shown in figure 157. The data revealthe following:
(I) The compositedata show little tendencyfor motion sicknessfor allturbulenceconditionsfor both SMCS off and on modes. FTE were more affected
than pilots.
: One subjectcommentedthat: 'Turbulenceis not a major influence. The rough-ness of terrainhas more impactwhen you are stuffedin that 'blackhole' withno windows." (FTEC)
227
PhysicalDiscomfort
All subjectswere asked to rate the degree to which TF flight introducedany specialtendenciestoward physicaldiscomfort. Ratingswere requested:forthe conditionsof SMCS off and on for four levelsof turbulence: smooth
air and light,moderate,and heavy turbulence. The combinedratingsfor allcrewmen are shown in figure 158. The data'revealthe following: -"
(I) Physicaldiscomfortincreasesas turbulenceincreases;althoughatheavy turbulence,the effectsare rated only as slight.
(2) Less physicaldiscomfortis reported for light,moderate,and heavyturbulencewith SMCS on than with SMCS off. For smooth air, there is nodifferencein physicaldiscomfortratingsfor the SMCS off versus SMCS onmodes.
One subjectcommentedthat: "Physicaldiscomfortresults from feelingsofirritation,aggravation,and anxietyproducedby the rough ride and highworkload,hard to pin down further." (PLTD)
AdditionalRide QualityObservations
All subjectswere asked to add any additionalobservations(not coveredby the ride qualityquestions)concerningride qualitycharacteristicsorSMCS effectsduring TF flight or any other characteristicsrealtingto TFwhich have a bearingon crew comfortand efficiency. Comments includedthefollowing:
(i) "SMCS is very effectiveand required in the B-I to aid the flightcrew in performanceof the TF task." (PLTA)
(2) "Essentialfor effectiveB-I MrF, desiredfor effectiveB-I ATF."(PLT C)
L
(3) "Turbulencemakes this aircraft hard to stabilizeon a bank angle, --adds to an alreadyhigh workload." (PLT D)
(4) "Ride qualitieswith SMCS are definitelybetter than without - thedifferenceI don't think is reallyable to be seen in the layout of thisquestionnaire." (FTEA)
(5) "Lack of outsidevisual referencecontributessignificantlytodisorientationand motion sickness. Movementabout the crew compartmentishamperedby TF, especiallyin turbulence. Any movement in the crew compartment
228
also contributesto disorientationand motion sickness. There is a significantimprovementin the ride with SMCS on." (FTEB)
(6) "I considerSMCS essentialfor long term TF flight." (OSOA)
(7) "Due to part of the controlsand displayslocation,it is difficultto accomplishsome weapon-orientedtasks under turbulentconditions. Thesensitivityof both the navigation (NAV)panel and Stores ManagementSystem(SMS)FWD/REVswitchmake it difficultto use under turbulentflight condi-tions." (OSOB)
(8) 'The discomfortsin the OSO stationare are severein light ormoderate turbulencewith SMCS off due to the fact that we have no outside
reference." (OSOC)
HandlingQualities
Pilots were asked to describethe effects,if any, of SMCS activationonaircrafthandlingqualitiesduringMTF. Commentsincludedthe following:
Pitch Control
(I) "A/C ride is smootherwith SMCS on and thereforecontrolis easier."(PLT A)
(2) "Assists by damping." (PLT B)
(5) "COUld have some detrimental effect because rapid control inputsare apparently countered by SMCS -- not considered a problem to date, moreevaluation required." (PLT C)
(4) "Increases pitch forces, se_ns to slow aircraft response." (PLT D)
Lateral/Direct ional Control
(1) "A/C ride is moother with SMCSon and therefore control is easier."(PLT A)
(2) "Assists by damping." (PLT B)
229
APPENDIX
NONENCIATURE
This reportwas the resultof the contributionsof a number of authorsand each has used nomenclatureunique to his particulardiscipline. In order -_to help the reader to quicklylocate a given symbol,this sectionhas beenorganizedso that a general section,is presentedfirst followedby nomenclatureassociatedwith three sectionsof this reportwhich are especiallyheavy inspecializednomenclature. These sectionsare: "FlexibleAircraft Equationso[ Motion," "Impactof SMCS on SelectedLoads,"and "_S Vane Effect on Inlet/Engine Characteristics."Under this system,similarsymbolsoften have differ-ent meanings;the reader is cautionedto identifysymbolswithin the contextoftheir use.
GENERAL
A amplitudesettingof B-I oscillatingsystem for the SMCS vanes
A/C aircraft
alt altitude
ATF automatic terrain following
AUTO automatic
BP butt plane
b[_F referencelength
mean aerodynamicchord
CITS centralintegratedtest system
CG center of gravity
cm centimeters
deg degrees
El bending stiffness
[ frequency,cycles per second
F, Flex subscriptdenoting flexible
230
F1t flight
ft feet
];l'Ji flight test engineer
[:/R flexible-to-rigidratio
[
]R flexible-to-rigidratio of bracketedparameter[
FS fuselage station
g Accelerationof gravity
CJ torsionalstiffness
GVT ground vibrationtest
crew sensitivityindex; subscriptZ denotesvertical axis,li() Y denoteslateralaxis
Hz hertz (cyclesper second)
hr hours
in. inches
_K _bREFk reducedfrequency,_ , V
o o
K SMCS gain
Khp SCAS gain scheduledwith altitude
kg kilogram
- KMNn yaw SCAS lateralaccelerometergainY
kq pitch SCAS gym gain
k_ SMCS gain scheduledwith dynamicpressureq
231
K_ r yaw SCAS gyro gain
Knz pitch SCAS normal accelerationgain
ib pounds
L-gain lateralSMCS gain
Lh dimensional,force along Z-axisdue to plungingmotion h,+ down
Le , dimensionalforce along Z-axisdue to pitchingmotion o,+ downJ
[ALO]Real [L@I Real V° [Lh]Imag
Lni dimensionalforce along Z-axis due to structuralmodegeneralizedcoordinatemotion hi, + down
Lwg dimensionalforce along Z-axisdue to verticalgust velocityVCg,+ down
L_ dimensionalforce along Z-axis due to control surfacemotion _,+ down
m meter
M Mach number
MrF manual terrainfollowing
Mh dimensionalmoment about Y-axis .dueto plungingmotion h,+ nose up
Mo dimensionalmoment about Y-axis due to pitchingmotion 6,+ nose up
Mni dimensionalmoment about Y-axis due to structuralmode gen-eralizedcoordinatemotion hi, + nose up
232
Mwg dimensionalmoment about Y-axis due to verticalgust velocityw , + downg
N newton
N ntmber of test participants
Qih dimensionalgeneralizedforce in structuralmode i due to
plungingmotion h, + for ni increased
Qi@ dimensionalgeneralizedforce in structuralmode i due to
pitchingmotion @, + for Ni increased
[AQi_IRoal IQi0]Real-(_) [Qih]imag
Qini dimensionalgeneralizedforce in structuralmode i due tostructuralmode generalizedcoordinatemotion hi, + for niincreased
Qiwg dimensionalgeneralizedforce in structuralmode i due toverticalgust velocityWg, + for Ni increased
Qi6 dimensionalgeneralizedforce in structuralmode i due tocontrol-surfacemotion _, + for ni increased
SL sea level
V-gain vertical SMCS gain
WL waterline
dampingratio
@ pitch angle about elastic axis
A sweep angle of leading edge of lifting surface
a Vertical gust intensityderived from angle-of-attackvane_Vmeasurements
OBv lateralgust intensityderivedfrom sideslipangle vanemeasurements
elasticaxis-bendingslope
233
•() power spectraldensityof subscriptparameter
frequency,radiansper second
RIDE QUALITYEQUATIONSOF MOTION RELATED
qo 1/2 pV2, dynamicpressure
p density of air
AP
_o pressurecoefficient
Vo resultantvelocity of the CG
componentof resultantvelocity;subscriptdenotesaxisV r ) along which componentacts
Wg verticalcomponentof gust velocity
Vg lateralcomponentof gust velocity
Sw wing area
bw wing span
_w wing mean aerodynamicchord
Z vertical deflection
Y side deflection
Zx' Zy' _z distancealong the x, y, and z-axis,respectively
- - distancefrom vehicle CG to control surfacek hingeline -_' z( ) (+ aft), subscriptidentifiessurface
distance from control surfacehingelineto surfaceCG_' _( ) (+ aft), subscriptidentifiessurface
distancefrom X-axis to surfaceCG in-planeperpendicularto) plane of symmetry (always+), subscriptidentifiessurface.
_°F perpendiculardistancefrom CG to thrust axis; + down
234
W airplane weight
m airplane mass
mass of control surface,subscriptidentifiessurfacem, m()
M. the ith mode generalizedmass,1
f f f dX®dzy,z) Am(x,y,
Ix, Iy, Iz moment of inertiaabout the X-, Y-, and Z-body axis, respectively
I product of inertia; positive when the principal X-axis isxzbelow the body axis at the nose of the vehicle
IR engine rotor moment of inertia
I( )[L moment of inertiaabout hingeline; ( ) subscriptidentifiessurface
n( ) load factor;subscriptdenotesaxis along whichcomponent acts
Euier azimuthangle
@ Euler pitch angle
Euler roll angle see figure 6
X, Y, Z body-axiscoordinates
Xe' Ye' Ze earth-axiscoordinates
11 altitude (+ up from sea level)
235
p rollingrate about X-body axis •
p rollingaccelerationabout X-bodyaxis
q pitchingrate about Y-body axissee figure 2
q pitchingaccelerationabout Y-body axis
r yawing rate about Z-body axis
r yawing accelerationabout Z-bodyaxis
_R rotationalrate of engine rotor relativeto airframe
_. natural frequencyof ith modei
forcingfrequency
angle of attack;angle betweentheprojectionof the resultantvelocityvector on the XZ-planeand the X-body
(referenceaxis) _ see figure 5|sideslipangle; angle betweenthe Iresultantvelocity vector and the Jplane of symmetry XZ
I" control-vane dihedral angle
6 con.trol-surfacedeflection;positive deflectionproduces
( ) positive force (+CN,+Cy)(seefigure 3); subscriptidentifiessurface
control-surfaceacceleration;positive in the sense that 6( )6( ) is positive;subscriptidentifiessurface ,.
6H rolling tail controldifferentialdeflection,+ deflectionproduces +Cz
g accelerationof gravity
gsi structuraldampingconstant,mode i
236
n. deflectionof the ith normalizedstructurali mode at normalizationpoint
n. rate of change of the ith mode at pQint ofi normalization
_. accelerationof the ith mode at point ofi normalization > see figure 4
¢.( ) the i normalizedmode shape; i.e., ratio ofi local deflectionto deflectionat normalizing
point (nondimensional); ( ) superscriptdenotes location
_! )' slope of the ith normalizedmode; ( )i superscriptdenoteslocation
T
_! ) fuselagetorsionalangle, ( ) superscriptdenotes location1
F force
Z aerodynamicforce in Z-direction
N aerodynamicnormal force (N = -Z)
L aerodynamiclift force (L_N for small _)
C aerodynamicchord force (C = -X)
D aerodynamlcdrag force (D_ C for small _)
X aerodynamicforce in X-direction _ see figure 3
Y aerodynamicforce in Y-direction
T thrust (T = X)
L aerodynamicrollingmoment about X-axis
M aerodynamicpitchingmoment about Y-axis
N aerodynamicyawing moment about Z-axis
237
Nwg normal force due to unit vertical gust velocity
Mwg pitchingmoment due to unit vertical gust velocity
Q[ generalized force in ith mode fffF(x,y,,) i(x,y,z)dxdydz ..
generalized force in structural mode i due unit verticalQiwg gust velocity
Yvg side force due to a unit lateral gust velocity
rollingmoment due to a unit lateral gust velocityLvg
Nvg yawing moment due to a unit lateralgust velocity
Nni yawing moment due to mode i deflection
Yni side force due to mode i deflection
M_i pitching moment due to mode i deflection
Nni yawing moment due to mode i deflection
Lni rollingmoment due to mode i deflection
Qij generalizedforce in mode i due to mode deflectionj
C chord-forcecoefficientCC = Swqo
DCC= --chord-force coefficientdue to control-surfacedeflection,
CC6( ) a6()subscriptidentifiessurface
N . normal-forcecoefficientCN = %%
OCNCNf Off normal-force curve slope
238
8CNnormal-forcecoefficientdue to downwashlag and
CN_ = _Cw verticalacceleration
8CN
CNq = -/q Cw_ normal-forcecoefficientdue to pitch rate
8CN
CN__ = a/Q_.2\lwl normal-forcecoefficientdue to pitch acceleration\4V2/
8CN
CNni = 8Ni normal-forCedeflectioncoefficientdue to structuralmode
8CN normal-force coefficient due to structral mode
CNni = 8_, deflection rate
8CN normal-forcecoefficientdue to controlsurface
CN6 = 86( deflection,subscriptidentifiessurface() )
8CN normal-forcecoefficientdue to controlsurface
CN6( =) On deflectionrate, subscriptidentifiessurface• ()
M
Cm = °w_w_=-o pitching-momentcoefficient
8CmCm_ = 8_ pitching;moment curve slope
239
_Cm pitching-momentcoefficientdue to downwashlagCmt_ = &I \ and verticalacceleration
3Cln __
Cmq = __qC-w_ pitching-momentcoefficientdue to pitch rateH I
-\2Vo!
aCm
Cmq = /_F.,2\ pitching-moment coefficient due to pitch acceleration
_Cm pitching-moment coefficient due to structuralCm_i = 3ni mode deflection
3Cm pitching-momentcoefficientdue to structuralmode
i _ deflectionrate
Cm pitching-momentcoefficientdue to control surface
Cm6( =) 36( ) deflection,subscript identifiessurface
8Cm pitchingmoment due to control surfacedeflectionrate, subscriptidentifiessurface
) )°
= a subscript_, B, etc, as shown indicatesthat()(_) bracketedparameter is a nonlinearfunctionof the
subscriptvariable
Qi generalized-forcecoefficientin the ithCni =
Swq° mode
240
C . = __8C_i_ generalized-forcecoefficientdue to angle of attack_i_ 8_ in ith mode
8Cni generalized-forcecoefficientdue to downwashlag
C_i& = 81_Cw_1-----'_ and vertical acceleration
VV°I
aCni generalized-forcecoefficientdue to pitch rate
C_iq = /qC-w_ in ith mode
3Cni generalized-forcecoefficientdue to pitch
Cni_l -tqcw21 acceleration in the ith mode
_CNi generalized-forcecoefficientdue to the jth mode
C_I_j" = _n]. shape in the ith mode
3CNi generalized-forcecoefficientdue to rate of change
C_i_J -- (-_o)
"" fli of jth mode in ith mode
aCni generalized-forcecoefficientdue to control-surfacedeflection in ith mode, subscript identifies control
CrJi_( ) a6( ) surface
aCNi generalized-forcecoefficientdue to control-surface
=) a6( ) k deflectionrate in ith mode, subscriptidentifiescontrolforce
YCy -- - side-force coefficient
qoSw
aCy
Cy_ = _s side-forcecoefficientdue to angle of sideslip
241
_Cy side-forcecoefficientdue to rate of change ofCy_i = O/_bw\ angle
12---_°_ sideslip (lateralacceleration)
CyBqi = 3CyB side-forcecoefficientdue to sideslipangle caused -_by symmetricstructuralbending (dihedral)i
_Cy
Cyr = side-forcecoefficientdue to yaw rate01 l\2Vo!
2%
CYr = a/_b,.2\,___I side-forcecoefficientdue to yaw acceleration\4Vo2 /
_Cy
Cyp = /pbw \ side-forcecoefficientdue to roll rate
_Cy
Cy_ = a/_bw2\| | side-forcecoefficientdue to roll acceleration\4Vo2/
_Cy side-forcecoefficientdue to control-surfaceCy6 =
( ) _6( ) deflection,subscriptidentifiessurface
_Cy side-forcecoefficientdue to control-surface
Cy_( = _() ) deflectionrate, subscriptidentifiessurface _
242
J
N
= Swbwqo yawing-momentcoefficientCn
_Cn
C_fl = _ yawing-momentcoefficientdue to sideslipangle
_Cn yawing-momentcoefficientdue to rate of changeCnfi =
_bw) of sideslipangle (lateralacceleration)
;CnB yawing-moment coefficient due to sideslip angle
Cn_ni _i caused by symmetricstructuralbending (dihedral)
_Cn
Cnr = /rbw\ yawing-momentcoefficientdue to yaw rateI I
_CnCn. = yawing-momentcoefficientdue to yaw acceleration
r i_bw2_
_Cn
Cnp = 'Jpbw\lI yawing-momentcoefficientdue to roll rate\2Vo/
_Cn
Cn_ = D/pbw2_ yawing-momentcoefficientdue to roll acceleration
\4Vo J.
_Cn= yawing-momentcoefficientdue to control surface
Cns( ) 26( ) deflection,subscriptidentifiessurface
243
aCn yawing-momentcoefficient due to control surface(:n_ =
( ) a_( ) deflection rate, subscript identifies surface
LCI = rolling moment
Swbwqo --
aCl
Cz_ = aB rolling-momentcoefficientdue to sideslipangle
aC_ rolling-momentcoefficientdue to rate of change --C_ =
81_bwl of sideslipangle (lateralacceleration)\ -vo!
aCz8 rolling-momentcoefficientdue to sideslipangle
C_Bni = ____aNi caused by symmetricstructuralbending (dihedral)
aC_
C_r = /rbw\ rolling-momentcoefficientdue to yaw rate
_C_
C_r = /+h2\ rolling-momentcoefficientdue to yaw acceleration
aC
Cgp = o/pb\|___Z_.lrolling-momentcoefficientdue to roll rate
ac= _
Cg_ /nh2\ rolling-momentcoefficientdue to roll acceleration
i wj8\4V02/
244
aC£ rolling-momentcoefficientdue to control surface
C_6( ) a6( ) deflection,subscriptidentifiessurface
aC_ rolling-momentcoefficientdue to control surface
_ C_( ) an( ) deflectionrate, subscriptidentifiessurface
aCNi generalized-forcecoefficientdue to sideslipangleC_i = a6 in ith mode
8
_CNi generalized-forcecoefficientdue to sideslipangleCni" =
6 l_bw_ rate of change in ith mode
aC.ni= 6 generalized-forcecoefficientdue to sideslipin the
Cni6 an. ith antisymmetricmode due tobending in the jthnj J symmetricmode
aCni generalized-forcecoefficientdue to yaw rate in
Cnir =Irbw_ ith mode
aC_i generalized-forcecoefficientdue to yaw accelerationCni" = • 2_ in the ith mode
r rbw
_ aCni generalized-forcecoefficientdue to roll rate in
CHip =pbwl ith mode
\2Vo!
_C_i generalized-forcecoefficientdue to roll acceleration
Cni_ = i_bw2 _ in the ith mode
a t[ 14Vo2 I
245
qoSwKs =
mVo-
qoSw_wKq = ly
qoSwK# = mVo
qoSwbwKr =
Iz
qoSwhwKp =
Ix
qoSwKni = Mi
i /1-
i, j used as subscriptto identifystructuralmode
k subscriptidentifyingcontrol surface
II subscript identifying horizontal-tail control surface
r subscriptidentifyingrudder
cv subscriptidentifyingstructuraln_de controlvaner
R subscriptindicatingreal part
I subscriptindicatingimaginarypart
o subscript indicating trim value
tL hingeline
I;RL fuselagereferenceline
246
MAC mean aerodynamicchord
( )TF aerodynamictransferfunction
- LOAD EQUATIONSOF MOTION RELATED
a accelerationl
A gust response factor Oo/Ow
b turbulencefield parameterdenotinggust intensity
IB ] phased loadingconditions,one column per condition,each
elementproportionalto Pi-3 °" _"i 3
IE J expectedload values for each load it_n
{F} forces at each SIC point: real or in the frequencydomain,complex
g structuraldampingparameter
Ih I real generalizedcoordinates
H(_) frequencyresponsefunctionfor a load item
II*(_) denotes complexconjugateof H(_)
Ill(w)] frequencyresponsefunctionsof load items, one row per itemfor i01 frequencies
j #T
- [LOADS ] loads; shears,moments, and torques;real or in the frequencydomain,complex
LOAD geometryto compute shears,moments,and torques,one rowGEOM per load item, one column for each SIC load point
IDAD I load generationmatrix, one row per load item, one columnGEN 3 for each normal elasticmode
EM_ mass matrix
247
[_M] generalizedmass matrix for rigid-bodymotions
[M_ generalizedmass matrix = [M][_],rigid and elasticmodes -_[¢]T
\
N number of crossings of zero per hour with a positive slopeo
N(y) averagenumber of level crossingswith a positive slopeequallyor exceedingy per hour
P turbulencefield parameterdenoting proportionof time inturbulence
q I complexgeneralizedcoordinates
QM] generalizedaerodynamicforces due to modal motion
Qc] genezalizedaerodynamicforces due to unit deflectionsofcontrolsurface,column 1 for horizontaltail and c61umn 2for the mode controlvane
_iI generalizedforces due to a unit sinusoidalgust
R] rigid-bodymode-modifyingmatrix
S Laplace operator
S_ phased loadingconditions,one column per condition,eachcolumn scaled so that the diagonal elementequals the corre- .-'-spondingexpectedload value
[ Sl C] structural flexibility influence coefficients
ISLOADS I phased loading column from [S],shears,moments, and torques
IT .] transformationmatrix,for controlsystem feedback,relatingcontrol-surfacedeflectionsto the generalizedcoordinates
V velocity
248
_H' _cv control-surfacedeflections;subscriptH for horizontalstabilizer,subscriptV for mode controlvane
mode deflection
L[ ¢]J mode shapesby columns, superscriptdenotestype of modes;RBM for rigid-body,and E for normal elastic
_w' _o root mean-squarevalues of gust velocity and output itemresponse,respectively
oi, o. root mean squarevalue of responseitems i and jJ
_(_) gust power spectrum,normalizedon unit gust intensity,o2W
(_) load item output responsepower spectrumo
pitch rate
o.. correlationcoefficientexpressingthe degreeof statisticalij linear dependencebetween load item i and load item j
frequency
structuralnormal elastic-modefrequency
] squareor rectangularmatrix;where i = number of rows, andi,j j = number of colunns
E _ diagonalmatrix
cQlumn matrix
,_. L _] row matrix
[_T matrix transpose
[] matrix inverse
249
ENGINE/INLET RELATED
A/l) analog to digital
AIP inlet/engine aerodynamic interface plane
AIS air induction system
Aduct duct area, psi/count
CAL calibration
CBW constant bandwidth
CTS counts
CTSCAL counts output during calibration step
CTS_4EAN mean value in counts during operate step
CTSOP counts output during operate step
CTSZ counts output during zero step
db decibels
I!CI-8 inlet configuration identification
FS full scale
HP high pass
IDC circumferential distortion component _.
IDL engine stall-margin ratio (function of IDC and IDR)
IDR radial distortion component
ips inches per second
K i000 feet
LP low pass
25O
NF engine fan speed
N2 nitrogen
P pressure
_P pressure difference
PCM pulse-codemodulation
PDYN Instantaneouspressure (dynamic)
PLA engine power lever angle
PPCM time-averagedpressure (steadystate)
pps pounds per second
psi pound per square inch
PT total pressure
PTI average total pressureat inlet/engineAIP
VrI local total pressureat inlet/engineAIP
PTO free-streamtotal pressure
FIIv_X maximum total pressure at inlet/engineAIP
VFMIN minimum total pressure at inlet/engineAIP
RB first movable inlet ramp angle
RC secondmovable inlet ramp angle
R2 referencepressure
SMCV structuralmode controlsystemvane
sps sampleper second
VCO voltagecontrolledoscillator
WIR correctedengine airflow
251
X2 amplifier (times 2)
X6 amplifier (times 6)
ZOC a three-way value (refer to page 161)
t_
Z_ fractional increment on total pressurecontour, AIP
6002 reference pressure, absolute units
6631 calibration pressure, absolute units -_
252
REFERENCES
I. Wykes, John H.; Borland,ChristopherJ.; Klepl,Martin J.';and MacMiller,
_ Cary, J.- Design and Developmentof a StructuralMode Control System,NASA CR-143846,October1977.
2. Etkin, Bernard: Dynamicsof Flight. John Wiley and Sons, Inc., New- York.
3. Bisplinghoff,RaymondL.; Ashley,Holt; Halfman, RobertL." Aeroelasticity.Addison-WesleyPublishingCompany, Inc., Reading,Massachusetts.
4. Landahl,M. T.: GraphicalTechniquesfor AnalyzingMarginallyStableSystems. Journalof Aircraft,September-October1964.
5. Rustenburg,John W.: Developmentof TrackingError FrequencyResponseFunctionsand Aircraft Ride QualityDesign Criteriafor Vertical andLateralVibration. ASD-TR-70-18,January1971.
6. Stenton,Thomas E.: TheoreticalFrequencyResponseFunctionsand PowerSpectraof the XB-70 Responseto AtmosphericTurbulence. NASA CR-1621,August 1970.
7. Dugundji, John: On the Calculationof NaturalModes of Free Free
Structure. Journalof the AeronauticalSciences,February1961.
8. Flight ControlSystem Description,Rockwell International,North American
Aircraft Division,E1 Segundo,TFD-71-807_June 1971 (RevisionAugust19753.
9. Austin, WilliamH., Jr.: Developmentof ImprovedGust Load CriteriaforUnited StatesAir Force Aircraft. SEG-TR-67-28,September1967.
i0. Fuller,J. R.; Richmond,L. D.; Larkins,C. D.; and Russell,S, W.:Contributionsto the Developmentof a Power SpectralGust Design Procedurefor Civil Aircraft. FAA-ADS-54,January1966.
253
J
1. ReportNo. 2. GovernmentAccessionNo. 3. Recipient'sCatalogNo.NASA CR-144887
4. Title and Subtitle 5. Report Date
January. 1980ANALYSES AND TESTS OF THE B-1 AIRCRAFT 6. PerformingOrganizationCodeSTRUCTURAL MODE CONTROL SYSTEM
7. Author(s) 8. PerformingOrganizationReportNo.John H. Wykes, Thomas R. Byar, Cary J. MacMiller, and NA-79-405David C. Greek
10. WorkUnitNo.9. PerformingOrganizationNameandAddress
Rockwell InternationalNorth American Aircraft Division 11. Contractor GrantNo.
815 Lapham Street NAS4-2519El Segundo, CA 90245 "
13. Typeof ReportandPeriodCovered12. SponsoringAgencyNameandAddress
Contractor Report - Final
National Aeronautics and Space Administration 14. SponsoringAgencyCodeWashington, D.C. 20546
H-1109
15. SupplementaryNotes
NASA Technical Monitors: Jim MeKay and Larry Felt, Dryden Flight Research CenterNASA Program Manager: Jack Nugent, Dryden Flight Research Center
16. Abstract
An 18-month program was conducted to compile and document for publicationinformation pertaining to analyses and flight tests of the B-1 Structural ModeControl System (SMCS). This is the second phase of a c_ntinuing effort;results from the first phase study are documented in De_;ign and Developmentof a Structural Mode Control System, by John H. Wykes, Christopher J. Borland,Martin J. Klepl, and Cary J. MaeMiller (NASA CR-143846, October 1977). This
report covers the_following topics: _'_
(1) Flexible aircraft equations of motion
(2) Description of flexible aircraft analyses model
(3) Comparison of analyses and flight-test performance results of the SMCS
(4) A summary of the study of the forward SMCS sensor package relocation
(5) Truncated analytical models used in simulation effort
(6) An analysis of the SMCS vane interference effects
(7) Impact of SMCS on selected loads
(8) Flight-test results of the SMCS vane effects on inlet/engine characteristics
r (9) Summary of SMCS flight-test results
17. Key Words(SuggestedbyAuthor(s)) 18. DistributionStatement
Structural mode control systemLarge flexible aircraftAnalyses and tests
STAR category: 05
19. SecurityClassif.(ofthisreport) 20. SecurityClassif.(ofthispage} 21. No.of Pages 22. Price*Unclassified Unclassified 268
*For sale by the National Technical Information Service, Springfield, Virginia 22151
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