TOFinal Report for Period September 1980 -February 1982 SUBJECT TO EXPORT CONTROL LAWS This document...
126
UNCLASSIFIED AD NUMBER ADB069403 NEW LIMITATION CHANGE TO Approved for public release, distribution unlimited FROM Distribution authorized to U.S. Gov't. agencies only; Critical Technology; Test and Evaluation; FEB 1982. Other requests shall be referred to Air Force Aero Propulsion Laboratory [AFWAL/POTX], Wright-Patterson AFB, OH 45433. AUTHORITY 27 Jun 1990, ST-A per WRDC-IST notice THIS PAGE IS UNCLASSIFIED
Transcript of TOFinal Report for Period September 1980 -February 1982 SUBJECT TO EXPORT CONTROL LAWS This document...
UNCLASSIFIED
AD NUMBER
ADB069403
NEW LIMITATION CHANGE
TOApproved for public release, distributionunlimited
FROMDistribution authorized to U.S. Gov't.agencies only; Critical Technology; Testand Evaluation; FEB 1982. Other requestsshall be referred to Air Force AeroPropulsion Laboratory [AFWAL/POTX],Wright-Patterson AFB, OH 45433.
AUTHORITY
27 Jun 1990, ST-A per WRDC-IST notice
THIS PAGE IS UNCLASSIFIED
AFWAL-TR- 82-2017
C TRANSONIC FAN/COMPRESSOR'Q ROTOR DESIGN STUDY
Volume III
D.E. Parker and M.R. SimonsonCZ) General Electric Company
Aircraft Engine Business GroupAdvanced Technology Programs Dept.Cincinnati, Ohio 45215
February 1982
Final Report for Period September 1980 - February 1982
SUBJECT TO EXPORT CONTROL LAWSThis document contains information for manufacturing or using mu.nitions of war Export of the information contained herein, or releaseto foreign nationals within the United States, without first obtainingan export license, is a violation of the International Tratfictin.ArmsRegulations. Such violation is subject to a penalty of up to 2 years im.prisonment and a fine of $100,000 under 22 USC 2778.
Include this notice with any reproduced portion of this document.
Distribution limited to U.S. Government gencies only: Testand Evaluation, February 1552. O1W requests tor Ohfs dowu-nent must be reerred to Air Forve Aero Propulsion Labomory(A"WAL/POTX), Wright-Pattorson AFS, ON 45433
DEC 1 1982Aero Propulsion Laboratory A
LLJ Air Force Wright Aeronautical Laboratories- - Air Force System Command
Wright-Patterson Air Force Base, Ohio 45433
82 12 01 024
NOTICE
When Government drawings, specifiations, or other data are used for anypurpose other than in conjunction with a definitely related Government procure-ment operation, the United States Government thereby incurs no responsibilitynor any obligation whatsoever; and the fact that the Government may have formu-lated, furnished, or in any way supplied the said drawings, specifications, orother data, is not to be regarded by implication or otherwise as in any mannerlicensing the holder or any other persons or corporation, or conveying anyrights or permission to manufacture, use, or sell any patented invention that
!1 may in any way be related thereto.
This technical report has been reviewed and is approved for publication.
ARTHUR J. WENNERSTROM WALKER H. MITCRELLChief, Compressor Research Group Chief, Technology Branch
FOR THE COMMANDER
H. WAN BIDirector, Turbine Engine Division
"1
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AFWAL/POTX, WPAFB OH 45433 to help us maintain a current mailing list.
Copies of this report should not be returned unless return is required bysecurity considerations, contractual obligations, or notice on a specificdocument.
1
UnclassifiedSECURITY CLASSIFICATION 0F THIS PAGE (1ten Det, Entered)
T IO PG READ INSTRUCTIONSREPORT DOCUMENTATION PAGE BEFORE COMPLETING FORM,
1. REPORT NUMBER 2. GOVY ACCESSION NO. 3. RECIPIENT'S CATALOG NUMBER
AFWAL-TR-82-2017 I - 0 /_
4. TITL.E (a,.nd $ubti) S. TYPE OF REPORT & PERIOD COVERED
Transonic Fan/Compressor Rotor Design Study Final Technical ReportVolume III September 1980 - July 1982
6. PERFORMING ORG. REPORT NUMBER
R82AEB328, Volume III7. AUTHOR(@) S. CONTRACT OR GRANT NUMBER(@)
D.E. Parker F33615-80-C-2059M.R. Simonson
S. PERFORMING ORGANIZATION NAME AND ADDRESS 10. PP)GRAM ELEMENT. PROJECT, TASK
General Electric Cbmpany ANEA 6 WORK UNIT NUNBERS '
Aircraft Engine Business Group Project 2307
Cincinnati, Ohio 45215 Task SI____________________________________ Work Unit -. 38
11. CONTROLLING OFFICE NAME AND ADDRESS REPAero Propulsion Laboratory (AFWAL/POTX) u. R jbe'r 1982Air Force Wright Aeronautical Laboratories s. UMBER Of PAGESAir Force System CommandWright-Patterson Air Force Base. Ohio 45433 Ila
14. MONITORING AGENCY NAME 6 AODRESS(II/ll.,ent froem ConttolinS Office) IS. SECURITY CLASS. (of thl report)
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IS. DISTRIBUTION STATEMENT (of thle Report)
Distribution limited to U.S. Government agencies only: Test and Evaluation," dpLberl19$Sr Other requests for this document must be referred to AirForce Aero Propulsion Laboratory (AFWAL/POTX), Wright-Patterson AFB,
OH 45433.
17. DISTRIBUTION STATEMENT (of the Pbst,.ct entered In Block 20, Ifdillerent from Report)
IS. SUPPLEMENTARY NOTES
19. KEY WORDS (Continue on revere. old. $1 nocoeoary and identify by block nuvb.,)
Fan Aircraft Engines
Compressor Blade Thickne)s
Rotor Camber Distribution
Aerodesign Throat MarginAerodynamics
SO. ABSTRACT (Continue an reverse side It necoeeay and Identify by block number)
Volumes I through VI of this report describes the aerodynamic design
cf a series of five transonic rotors all parametrically related to a base-
line design documented in Technical Report AFAPL-TR-79-2078' Each of thefive designs deviate from the base line, in so far as practical, by a
variation of one parameter only. The parametric variations are specified
at the rotor tip. The original hub characteristics were preserved to the
maximum extent practical. The varied parameter was adjusted along the span.
DDI JAN73 1473 EDITION OF NOV65 OBSOLETE Unclassified
SECURITY CLASSIFICATION OF THIS PAGE (ften Date Entered)
UnclassifiedSECURITY CLASSIFICATION OF THIS PAGE(W1,w' Date Entetd)
:This volume describes the aerodynamic design details of the Phase IIrotor. The Phase II rotor has the tip airfoil maximum thickness locatedat 55% of meanline length as compared with 70% for the baseline rotor.The location of maximum thickness varied linearly with stream functionto 56% of meanline length at the hub, which is the same as the baselinerotor.
Unclassified
SECURITY CLASSIFICATION OF THIS PAGE(When Date Entoed)
VOLUME III
PHASE II ROTOR DESIGN
Foreword
This Final Technical Report was prepared by the Advanced TechnologyPrograms Department, Aircraft Engine Business Group, General Electric Company,Evendale , Ohio for the United States Air Force Systems Command, Air ForceWright Aeronautical Laboratories Wright-Patterson Air Force Ba_!-, Ohio underContract F33615-80-C-2059. The work was performed over a period ofone year starting in September 1980. Effren Strain (Captain USAF) was theAir Force Project Engineer for this program.
This report describes the results of an effort to aerodynamically definefive rotor designs, all parametrically related to a base line design whichcould b evaluated by future testing in order to define the sensitivityof transonic blade rows to several design variables.
For the General Electric Company Mr. D.E. Parker was the TechnicalProgram Manager for this program. Mr. M.R. Simonson was the principalinvestigator. Mr. A.J. Bilhardt was the overall Program Manager.
It
--- j
tii
VOLUME III
TABLE OF CONTENTS
SECTION PAGE
X. DESIGN OF PHASE II ROTOR 1
1. Introduction 1
2. Design Procedure 1
Xi. DETAILS OF PHASE II ROTOR DESIGN 11
1. Circumferential Average Flow Solution 11
2. Streamsurface Blade Coordinates 41
3. Plane Section Blade Coordinates 61
XII. REFERENCES 116
4-.
VOLUME III
LIST OF ILLUSTRATIONS
FIGURE PAGE
32. Phase II Rotor Static Pressure Distribution 3
33. Phase II Rotor Intrablade Work Distribution 4
34. Phase II Rotor Incidence Angle Versus FractionalImmersion 6
35. Phase II Rotor Deviation Angle Versus FractionalImmersion 6
36. Phase II Rotor Intrablade Departure AngleDistribution 7
37. Phase II Rotor Streamsurface Tip Section Comparedwith Baseline Design 8
38. Phase II Rotor Deviation Angle Minus ReferenceDeviation Angle Compared with Data Match 9
39. Phase II Stator Incidence Angle Compared withData Match 9
40. Phase II Rotor Throat Margin Compared withData Match 10
41. Compressor Flowpath with Calculation Stations 12
42. Stacked Phase II Rotor Streamsurface Sections 42
43. Stacked Phase II Rotor Plane Sections 43
vi
LIST OF SYKZOLS AND ABBREVIATIONS
1. Used in Circumferential Average Flow Output Tables
STA calculation station number
WTF total airflow
PSIC stream function (0 = tip (OD), 1 = hub (ID))
axial locatioa,. inches
R radius inches
PHI streamline slope degrees
CURV streamline curvature . = neg.,. - pos. 1/inches
VM meridional velocity ft/sec
CU absolute tangential velocity ft/sec
ALPHAM absolute flow angle on stream surface degrees
M meridional Mach number
SL calculation streamline number
BLDBLK flow blockage factor gree area - blocked are4/free area
PS static pressure psia
PT total pressure psia
TT total temperature degrees
BETAM relative flow angle on stream surface degrees
UREL relative velocity ft/sec
MREL relative Mach number
VABS absolute velocity ft/sec
MABS absolute Mach number
GAMA specific heat ratio
PT-RAT total pressure/inlet total pressure
TT-RAT total temperature/inlet total temperature
RCU radius x tangential velocity in-ft/sec
Ca axial velocity ft/sec
PCT IMM percent annulus i.ersion from tip (OD)
RAD average of leading and trailing edge inchesstreamline radii
ACC PTRATIO cumulative total pressure ratio
ACC TTRATIO cumulative total temperature ratio
vii
LIST OF SYMBOLS AND ABBREVIATIONS
1. Used in Circumferential Average Flow Output Tables (Cont'd)
AD. adiabatic efficiency
POLY polytropic efficiency
Axial axial velocity ratio across blade rowVEL R
2. Used in Stream Surface Blade Coordinate Tables
PT point number
PCT X fraction of meridional distance from leading edge
X meridional coordinate on meanline inches
Y tangential coordinate on meanline inches
B*M meanline angle on stream surface degrees
T(M) thickness of blade perpendicular to meanline inches
XS meridional coordinate on suction surface inches
YS tangential coordinate on suction surface inches
XP meridional coordinate on pressure surface inches
YP tangential coordinate on pressure surface inches
3. Used in Plane Section Coordinate Tables
axial coordinate of stacking axis inches
R radius of coordinate system origin inches
MU tilt angle in axial direction degrees
ETA tlit angle in tangential direction degrees
RHO section height inches
PT point number
ALPHA axial coordinate inches
3ETA* meanline angle from axial degrees
UPSILON coordinate perpendicular to ALPHA and radius inches
PCT AL fraction of axial distance from leading edge
T/C local thickness/chord ratio
viii
SECTION X
DESIGN OF PHASE II ROTOR
I. INTRODUCTION
The specification of the chord-wise location of airfoil maximum thickness
of transonic/compressor rotors has often been defined more on the basis of his-
torical practice than on a knowledge of its aerodynamic effect. Research by
NACA in the 1950's generally indicated that as relative irlet Mach numbers rose,
it was desirable to %ove the location of maximum thicknass aft on an airfoil.
The early work, however, was done with airfoils having significant positive
camber. Today, many airfoils have little overall relative turning in the tip
region, and frequently have S-shaped mean lines: negative camber followed by
some positive camber. In some cases, a forward shift in maximum airfoil thick-
ness may help achieve the desired airfoil suction surface shape, with a less
S-shaped mean line. There is also incentive to move the maximum thickness
forward to make the blade more capable of withstanding a bird strike without
excessive damage.
To get more definitive aerodynamic data on the effect of the location of
airfoil maximum thickness, the Phase I blade has been designed with the maximum
thickness located at 40%; the Phase II blade with maximum thickness at 55%, to
compare with the base line rotor which has its tip maximum thickness at 70% of
mean line length.
2. DESIGN PROCEDURE
The "data match" circumferential average flow solution, which was previously
described in Volume I, was used as a starting point for thp design of the Phase II
rotor. The annulus blockage used in the internal blade calculation stations was
adjusted to be consistent with the forward shift of the airfoil maximum thick-
ness. The assumed chord-wise distribution of work was iteratively adjusted to
obtain a calculated chordwise distribution of static pressure similar to that
of the data match calculations of the baseline rotor. Also the blade meanline
departure angle (the difference between the air angle and the meanline angle)
were adjusted to maintain the same throat area, and flow induction capacity as
the baseline blade. To adjust for the increased blade blockage in the forward
half of the blade, and to better match the data match status pressure distribu-
tion in the hub, the hub contour internally within the rotor was modified1
slightly relative to the baseline rotor. The Phase II rotor hub is .012 inches
lower in radius at the 60 percent chord location and is faired into the baseline
contour near the edge locations.
After each modification to the chordwise work distribution and/or departure
angles, revised blade annulus blockage and blade lean angles were calculated and
input to the circumferential Average Flow Determination (CAFD) computer program
for the next iteration.
The rotor exit radial distribution of total pressure and temperature was
maintained the same as the data match of the baseline rotor.
The resulting streamline static pressure distribution for the Phase II blade
is compared with the data match of the baseline rotor on Figure 32.
The assumed streamline work input (as a fraction of the total streamline
work) is plotted versus percent axial projection in Figure 33. The tip stream-
line is the one on the left. Each subsequent streamline is indexed to the right
by the value of its stream function (fraction of the total flow from the tip).
The dotted lines are lines of constant percent axial projection.
A method of characteristics computer program was used to analyze the flow
in the cascade flow induction region for streamlines 3 and 6 to assure that the
rotor would achieve the design flow. For other streamlines, the difference be-
tween the suction surface angle and the "free flow" streamline angle was com-
pared with similar data from the data match calculations of the baseline rotor.
This then, was used as a guide in setting the suction surface angle in the flow
induction region.
Somewhat larger incidence angles were used to help maintain the same flow
induction capacity as the baseline rotor. Phase II blade incidence angles are
shown on Figure 34.
A modified version of Carter's Rule was used to calculate a reference de-
viation angle for the baseline rotor. This procedure converts the vector dia-
grams (from the data match calculations) to an equivalent two-dimensional set
of vectors which would produce the same circulation as the actual blade taking
into account the change in streamline radius and meriodional velocity. The
difference between the deviation angle implied by the data match calculations
and the reference deviation angle was then added to the reference deviation2
-- -- -- ~ ~ b N.. .... .t .. .~7 . --
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. .. .....
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angle calculated from the modified Carter's Rule for the Phase II blade.
Phase II Rotor deviation angles are shown on Figure 35. A plot of departure
angles for each streamsurface section is shown in Figure 36. Once the intra-
blade work distribution was chosen these departure angles were required to
satisfy the desired incidence angles, deviation angles, and passage area ratios.
The resulting streamsurface tip section of the Phase II rotor is compared to
that of the baseline.rotor in Figure 37. The "deviation angle minus reference
deviation angle" for the Phase II rotor was kept essentially the same as the
data match analysis although there are some small differences. Figure 38
shows the "delta deviation" compared to the data match of the baseline design.
If the performance of a new rotor design is to be accurately evaluated
by comparing overall stage performance with the baseliie design then it is im-
portant that the stator have nearly the same entering conditions in both cases.
Figure 39 shows a comparison of the Phase II stator incidence angles with the
data match base. As can be seen the differences are small.
Figure 40 shows the radial distribution of Phase II rotor throat margin
and compares it to the data match case. The throat margin for a streamsurface
blade section is defined here as the percent of excess throat area over and
above the minimum theoretical area required to pass the streamtube flow at a
throat Mach number of 1.0 and assuming a total pressure loss equivalent to a
nornmal shock at the upstream Mach number. In a rotor the effect of radius
change (between the leading edge and throat) on the relative total enthalpy and
pressure is included. As can be seen in Figure 40 the Phase II rotor throat
margin is nearly identical to that of the data match of the baseline design.
Details of the Phase II rotor design are given in Section VIII.
5
" U .. .... . . . ..... .. . .. ....... ..... ....... ... .. .... . . . .
Iz
o * .....
5 I 7 8 9 10
INCIDENCE ANGLE, DEGREES
Figure 3._4_' Phase II Rotor Incidence Angle VersusFractional I.erson
.. ..
- 0 5 10 s 20 25
DEVIATION ANGLE DEGREES
Figure 35. Phase II Rotor Deviatnon AngleVersus Fractional I erson
.56
......
.... ..... ..
... . ........
-a.-4
A.-
04.
.7.
LEADING EDGE
ROTATION
BASELINE oe
~~4'e o,/;' __PHtASE II
Figure 37 • Phase II Rotor Streamsurface Tip Section Comparedwith Baseline Design
8ttN
• ............ .... .......... ! . ..... .... .... ......... ....... .... .. . .... .I|cz
° L PHASE II0
o- DATA JATCH'
0H co
.~ ~ ---- -...... ... ......- ... ....... .i ... .., ...... ... .......... ... .. .
1.5 2 2.5 3.5 4 4.5
DELTA DEVIATION, DEGREES
Figure 38. Phase II Rotor Deviation Angle MinusReference Deviation Angle ComparedWith Data Match
"/ ..~~.... .. .... .... ....:I '
0
PHAS IIo .MTC
( DATA
-1 0 1 2 3 4 5
INCIDENCE ANGLE, DEGREES
Figure 39 Phase II Stator Incidence Angle ComparedWith Data Match
9
'-0---- PHASE II
-- 0----DATA MATCH
0 t
C; ...... ....
4 6 ,8 to 12 14 lTHROAT MARGIN, PERCENT
Figure 40. Phase II Rotor Throat Margin
Compared With Data Hatch
10
SECTION XI
DETAILS OF PHASE II ROTOR DESIGN
1. CIRCUMFERENTIAL AVERAGE FLOW SOLUTION
The following tabulation presents the detail results of the Phase II Rotor
circumferential average flow computation. Each page of the tabulation gives
results for one calculation station. Figure 41 shows the calculation station
locations within the flowpath. At each calculation station various aerodynamic
parameters are given on each of thirteen calculation streamlines. Also given
are several mass averaged station flow properties. The Phase II rotor blade
forces are included at the end of this tabulation,
.4
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2. STRFAMSURFACE BLADE COORDINATES
Figure 42 shows the stacked Phase II rotor streamsurface sections. Each
page of the following tabulation gives the coordinates for one of these sec-
tions. The streamline designation for these sections corresponds to the calcu-
lation streamlines of the circumferential average flow calculation. Stream-
line 1 is at the casing and streamline 13 is at the hub. Also given in the
tabulations are coordinates for the section meanline, the meanline angle, and
the section thickness at each point. Streamsurface section chord, camber angle,
and stagger angle arc also given. All dimensions in this tabulation are in
inches or degrees.
41
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3. PLANE SECTION BLADE COORDINATES
Figure 43 shows the stacked Phase I rotor plane sections. The following
tabulation gives the coordinates for these sections. These sections are spaced
one half inch apart, beginning at the tip height of 8.5 inches and progressing
inward to 2.5 inches. These are the same section locations as given for the
baseline rotor in Reference 1. Also included in the tabulation are coordinates
for the section meanline, the meanline angle, and the section percent thickness
at each point. Plane section chord, camber angle, and stagger angle are also
given. These coordinates are intended to represent the blade under hot running
conditions and do not include any corrections for blade untwist, meanline defor-
mation, centrifugal growth or thermal growth.
61
/X
ROTATION
Figure 43 Stacked Phase II iwtor Plane Sections
62
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SECTION X1I
REFERENCES
1. A.J. Wennerstrom, and W.A. Buzzell, Redesign of a Rotor for a 1500 ft/secTransonic, High-Through-Flow, Single-Stage Axial-Flow Compressor withLow Hub/Tip Ratio, Air Force Aero Propulsion Laboratory, Wright-PattersonAFB, Ohio 45433, AFAPL-TR-2078, September 1979.
2. George R. Frost, Richard M. Hearsey, Arthur J. Wennerstrom, A Computer Pro-gram for the Specification of Axial Compressor Airfoils, Aerospace ResearchLaboratories, Wright-Patterson Air Force Base, Ohio 45433, ARL 72-0171,
3. Richard M. Hearsey, A Revised Computer Program for Axial Compressor DesignVolume I, Aerospace Research Laboratories, Wright-Patterson Air Force Base,Ohio 45433, ARL TF 75-0001, January 1975.
4. Arthur J. Wennerstrom, Personal Communication to L.H. Smith of GeneralElectric Company, September 12, 1980.
116
I,
