REPORT NO. NADC-81293-60

>-0-0 (._)

F-14 ROTARY BALANCE TESTS FOR AN ANGLE-OF-ATTACK RANGE OF 0° TO 90°

Billy Barnhart BIHRLE APPLIED RESEARCH, INC.

Jericho, NY 11753

JANUARY 1983

FINAL REPORT ~·-AIRTASK NO. A03V-32D/ 001 B/2F41-400-000

Contract No. N62269-82-C-0233 •'L, 1.-· '' ,_:

APPROVED FOR PUBLIC RELEASE: DISTRIBUTION UNLIMITED.

rrepared for Aircraft and Crew Systems Technology Directorate (Code 6053)

NAVAL AIR DEVELOPMENT CENTER Warminster, PA 18974

f tL •

.NOTICES

REPORT NUMBERING SYSTEM - The numbering of technical project reports issued by theNaval Air Development Center is arranged for specific identification purposes. Eachnumber consists of the Center acronym, the calendar year in which the number wasassigned, the sequence number of the report within the specific calendar year, andthe official 2-digit correspondence code of the Command Office or the Functional

* .Directorate responsible for the report. For example: Report No. NADC-78015-20indicates the fifteenth Center report for the year 1978, and prepared by the SystemsDirectorate. The numerical codes are as follows:

CODE OFFICE OR DIRECTORATE

00 Commander, Naval Air Development Center01 Technical Director, Naval Air Development Center02 Comptroller10 Directorate Command Projects20 Systems Directorate30 Sensors F Avionics Technology Directorate40 Communication & Navigation Technology DirectorateSO Software Computer Directorate60 Aircraft & Crew Systems Technology Directorate70 Planning Assessment Resources80 Engineering Support Group

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T. -. . -0 /7i

UNCLASS IFI EDCCuITY CLASSIFICATION OF THIS PAGE tRhm Doe #w64________________

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NAD-81 293-60 I~A/t ~4. TITLE (" S&ahtt) L S TYPE OF REPORT A PERIOD COVERED

F-14 ROTARY BALANCE TESTS FOR AN ANGLE-OF-ATTACK FinalRANGE OF 09"'TO 90?', S. PERFORMING ORG. REPORT NUMBER

7. AUTNORtO) 6. CONTRACT OR GRANT NUMS1911e)

Billy Barnhart N62269-82-C-0233

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Bihrle Applied Research, Inc.400 Jericho Turnpike AIRTASK NO. AO3V-32D/O0lB/Jericho. NY 11753 2F4l -400-non

11. CONTROLLING OFFICE NAME AND ADDRESS IS. REPORT DATEAircraft & Crew Systems Technology Directorate January 1983Naval Air Development Center (Code 6053) 1S. NUMBER OF PAGESWarmuinster, PA 18974 22

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UnclassifiedI". DEC ASFICATION/DOWNGRADIMG

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Approved for Public Release; Distribution Unlimited.

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IS. SUPPLEMENTARY NOTES

it. KEy WORDS eCelmhe mnvme old& it meemn OWdms by 6100k smhe)

F-14High angle-of-attack wind tunnel data

* Rotary BalanceSpinning

20. ABSTRACT (C00000 M veVW80 iWd i umminp7 UNI Idt OF 0100 hdA 1/12-scale model of the F-14 was tested on the rotary balance located in theLangley Spin Tunnel. Data were obtained for the basic airplane in the maneuverconfiguration with various control settings at three wing sweeps. The datawere supplied to the Naval Air Development Center on magnetic tape. Thisreport presents a description of these tests and the information supplied onthe data tape, as well as a list oll',spin modes predicted for the F-14 utilizingthe rotary balance data.<

DI JA 72DTO O 6'I USLE UNCLASSIFIEDShI016.L.01440 I SECUITY CLASSIFICATION OF THIIS PAME (WIM DmmLw

,,

NADC-81293-60

TABLE OF CONTENTS

PAGE

INTRODUCTION AccSNTIS GRA&I

SYMBOLS DTIC TAB 1

TEST EQUIPMENT Justification_ [13TEST PROCEDURES By_ _ 4Distribution/

MODEL Availability Codes 5

TEST CONDITIONS ist Avail and/or 6

DATA PRESENTATION 6

PREDICTED SPIN MODES 8

LIST OF TABLES

TABLE NO. PAGE

I. Dimensional Characteristics of the 1(12-Scale 10F-14 Model

II. Configurations Tested and Data File Index 11

III. Configuration Definition 13

IV. Comparison of Predicted and Experimental Spin Modes 14

V. Influence of Canards on Predicted Spin Mode 15

LIST OF FIGURES

FIGURE NO. PAGE

1 Photograph of 1/12-scale model installed on the 16rotary balance apparatus

2 Sketch of rotary balance apparatus 17

- 3 Three-view sketch of 1/12-scale model 19

4 Sketch of canards tested on 1/12-scale F-14 model 20

NADC-81293-60

j lmWRODUCT"ION

The Naval Air Development Center intends to conduct high angle of

attack studies on their Dynamic Flight Simulator modelling the

Navy/Grumnan P-14 airplane. As a part of this effort, rotary balance

wind tunnel force tests of an F-14 model were conducted to provide a

rotational aerodynamic data base up to 900 angle of attack.

A 1/12-scale model of the F-14 was tested on the rotary balance

located in the Langley Spin Tunnel. Data were obtained for the basic

airplane in the maneuver configuration with various control sottinss at

three wing sweeps. The data wore supplied to the Naval Air Development

Center on magnetic tape. This report presents a description of those

tests and the information supplied on the data tape, as well as a list of

spin modes predicted for the F-14 utilizing the rotary balance data.

0YOL5

The units for physical quantities used heroin are presented in U.S.

Customary Units.

b wing span, ft

a mean aerodynamic chord, ft

C ~ ~ ~~ xa azalfrc cefiietCA axial-fore oefficient, qS positive aft along

the body I-axis

drag coefficient,C. qS

-CN normal-force coefficient, positive upwardCN_ qS

e-1 -

. . . . .. . . . .*. . . -

· .. ·

c •

q

s

a

6 r

NADC-81293-60

froa the body X-Y plane

Side force aide-force coefficient, , positive out the qS

right wing

lift coefficient, Lift force qS

body-axis rollins-aoaent coefficient, Rollinl aoaent qSb

stability-axis rollina-aoaent coefficient,

Stab.-a, is roll aoaent qSb

pitchins-aoaent coefficient, Pitchina aoaent qSc

body-axis yawins-aoaent coefficient, Yawina-aoaent q~

stability-axis yawina-aoaent coefficient,

Stab,-axis yaw aoaept qSb

2 free-streaa dynaaic pressure, lb/ft

2 wing area, ft

ansle of attack, dog

ansle of ~ideslip, dea

angular velocity about spin axis, rad/sec

spin coefficient, positive for clockwise spin

differential horizontal tail deflection, positi,,e when risht surface is down, (6h -6h )/2, des

right left

s~etrical horizontal tail deflection, positive when trailins edge is down , des

rudder deflection, positive when trailing edge is to t he loft, deg

2

.. '

6 s p

NADC-81293-60

spoiler deflection, positive for riaht win& spoiler deflected trailins edae up

swoop analo of tho wins leadin& edae

Abbreviations cg center of aravity

RPM revolutions per •inute

trailing edge

TEST EQUIPMENT

A rotary balance •easures the forces and •omenta acting on a model

whi l e it is subjected to rotational flow conditions. A photograph and

sketch of the rotary ba l ance apparatus installed in the Langley Spin

Tunnel are shown in fia u res 1 a nd 2 , respectively. Tho syste•'s rotary

arm, which rotates about a ver ti cal axis at the tunnel center, is

supported by a horizontal boo• and is driven by a •otor kounted external

to the test section.

A NASA six-component strain aauge balance, affixed to the bottom of

the rotary balance apparatus and mounted inside the model, is used to

measure the normal, lateral, and lonaitudinal forces, and the yawina,

rollins. and pitching moments actina about the model body axis. Controls

located outside of the tunnel test section are used to activate motor s on

the rotary rig, which position tho model to the desired attitude. The

ansle-of-attack range of the rig is 0° to 90°, and the sidesl i p angle

range is ~15°. Spin radius and lateral displacement motors are used

t c position the moment center of the balance on, or at a specific

distance from, the spin axis. (This is done for each combination of angle

3

,:. .. ...... ... . . . . .. . . . . . . .

MADC-81293-60

of attack and sideslip angle.) It is customary to mount the balance to

the model such that its moment center is at the location about which the

aerodynamic moments are desired. Electrical currents from the balance and

to the motors on the rig are conducted through slip rings. Figure 2

illustrates various components of the rig and shows how the rig is

positioned in angle of attack and sideslip.

The system is capable of rotating up to 90 rpm in either direction.

A rane of b/l2V values san be obtained by adjusting rotational speed

and/or tunnel air flow velocity. (Static aerodynamic forces and moments

are obtained when 0-0.)

The data acquisition, reduction, and presentation system is composed

of a 12-channel scanner/voltmeter, a mini-computer with internal printer,

a plotter, and a CRT display. This equipment permits data to be

presented via on-line digital print-outs and/or graphical plots.

TEST PROCMES

Rotary aerodynamic data are obtained in two steps. First, the

inertial forces and moments (tares) sating on the model at different

attitudes and rotational speeds must be determined. Ideally, those

inertial terms would be obtained by rotating the model in a vacuum, thus

eliminating all aerodynamic forces end moments. As a practical approach.

this is approximated closely by enclosing the model in a sealed spherical

structure, which rotates with the model without touching it, such that

the air immediately surrounding the model is rotated with it. As the rig

4o-7

NADC-81293-60

is rotated at the desired attitude and rate, the inertial forces and

moments generated by the model are measured and stored on magnetic tape

for later use.

The second step is to remove the enclosure and record force and

moment data with the air on. The taros, measured in step one, are then

subtracted from these data, leasvin only the aerodynamic forces and

moments, which are converted to coefficient form and stored on magnotic

tape.

MODEL

A 1/12-scale model of the Navy/Grumman F-14 fighter airplane was

constructed of balsa and plywood. A three-view drawing of the model is

shown in figure 3, dimensional characteristics of the basic model are

listed in Table I , and a photograph of the model installed on the rotary

balance located in the Langley Spin Tuannel is presented in figure 1.

The model control surfaces could be set at any position prior to

testing. The maximum deflections for the control surfaces were:

rudder (deg) 30 right, 30 left

symmetrical horizontal tail (TE) 35 up, 10 down

differential horizontal tail (TI) 12 up. 12 down

spoilers (de) 55

Additionally, the slove vane could be extended to 150 or retracted

into the wing-glove and the maneuver flaps and slats could be extended

for the maneuver configuration to 10" and 70 deflections,

5

HADC-81293-60

respectively.

Canards were also constructed which could be mounted on the forward

fuselage sides, as shown in figure 4.

lS= COMMTONS

The tests were conducted in the spin tunnel at a free-stream

velocity of 25 ft/sec, which corresponds to a Reynolds number of

approximately 130,000 based on wing chord. All the configurations were

tested through an angle-of-attack range of 00 to 90° in $o

increments, unless otherwise noted in Table 11. For all the tests, the

spin azis passed through the full-scale airplane nominal cg location

(0.16ec). For each angle of attack, data were obtained at rotation rates

yielding the following Ilb/2V values, rounded to two decimal places:

0., 0.07, 0.14, 0.27. 0.41, 0.54. in both clockwise (positive) and

counter-clockwise directions.

DATA PRZBSTATON

Table 1'identifies the configurations tested and their

corresponding file names. The files are written in the order shown, on

a nine-track data tape. The data tape was produced by a CDC Computer and

should be read as a stranger tape with standard blocking at 1600 bpi. All

other parameters are the defaults. The data within each file is arranged

is the following order: The first 68 characters in each file provide

i"satification tuformatloa consisting of the date the tests were

6

...,.U-~t ~ * -~..

NAC-8129.-60

performed, the name of the model (F-14). the configuration tested, and a

run number used at the spin tunnel for bookkeeping purposes. These data

are written in A8.2A25,10.3 format.

- The remainder of each file contains the data measured during the

rotary balance testing of the given configuration, beginning with the

smallest angle of attack and most negative spin rate (Ob/ZY). The data

are ordered such that first the angle of attack is held constant while

the Jb/2V increases to its maximum. followed by the data for the next

larger angle of attack, beginning again at the most negative BbI2V. There

will be two sets of zero BbI2V data at each a, because static

(Bb/2V-0O) values were measured twice.

Each data point consists of 171 characters, which present the

following data in order:

angle of attack, sideslip angle, point number, CA. CY.

C . c . , n. /v,% . I ,CDCa'C C,- Ob/2VY CL- CD- C stab.'sttab.

Cns b, raw voltags for azial fores, side frtes.

normal force, roll moment, pitch moment, and yaw moment, and,

lastly, rotary rig RPM.

These data are written using the following format:

219.2,313.6,11F7.4,6F9.6,39.2

All the moment data are presented for a cg position of 0.160.

Unless otherwise designated in Table 11, the airplane was tested in

7

- . -.- .- - -. U ~ .. , . - . .. .. . . .. A

NADC-81293-60

the maneuver configuration, which consisted of the glove vans &ad maneuver

flaps and slats extended for the 220 and 50~ wing sweeps. At 638

wing sweep, the glove vane was extended.

Predieted Spin Nodes

A comparison of spin modes predicted using rotary balance data and

experimental. free-spinning model results are presented in Table IV. The

predicted spin node characteristics agree well with the experimental

results at 220 wing sweep. Both the predicted and experimental results

indicate that sweeping the wings to 6830 produces a slower, steeper spin

mode. although the rotary balance data shows a greater influence of wing

sweep than was observed in th. spin tunnel results. The effect of wing

* sweep is primarily caused by Improved damping of the yawing moment at

* high angles of attack as the wings are swept aft. This should also

provide improved recovery characteristics with the wings swept aft, as was

observed in the spin tunnel.

Spin modes with the stick aft (-23 0 ah ) are significantly

slower with pro-spin controls, and the results suggest that recoveries

can, likewise, be improved by maintaining aft stick. This is due to

improved yaw damping with the stick aft at high angles of attack.

When the airplane was configured with the canards shown in figure 4,

* steeper, slower spins resulted, as shown in Table V. This, again,

results from greater yaw damping at the high angles of attack, which.

in this case, is produced by the canards. The canards also eliminated a

non-zero yawing moment at zero sideslip angle and rotation rate in the

8

NADC-81293-60

400 to 70 angle-of-attack range. Therefore. the canards would be

expected to eliminate the 640 right spin predicted with aft stick and

neutral lateral controls, size* this mode is a result of the yawing

moet offset.

lo

NADC-81293-60

TABLE I.- DIMENSIONAL CHARACTERISTICS OF THE 1/12-SCALE F-14 MODEL

Overall length, ft . . . . . . ......... . . . . 5.16

Wing:Span (unswept), ft .*... .. .. .. ... 5.34Reference area, ft 2 . . . . o ...... 3.92Aspect ratio .... ....... 7 .28Taper ratio . . . . . . . . . . 0.265Sweep range, deg . . . 22 to 68Dihedral -lO50 'Root chord, in ........... ....... . . 13.93Tip chord, in ....... . . o . . . . . . . 3.69Mean geometric chord, in . . . . . ............ 9.80Incidence, deg:BL 8.025 o . . . . . .. . . . . . . +0.74BL 32.07 . ........ . . . . . . . . . . . -4.10

Pivot location FS 43.68BL 8.92

Horizontal tail:Span (overall), ft . ........ 2.73Area (exposed), ftz . . . ........... . . . . 0.97Aspect ratio (exposed), each .......... .. 2.56Taper ratio . . . ......... . . . 0.213Leading edge sweep, deg . . . . ........... 51Dihedral, deg . . ....... . . . . .. -3.5Root chord, in ........... ........ . . 12.38Tip chord, in .......... 2.64Mean geometric chord, in ...... 8.54Airfoil section:

Root . . . . . . . .. . . . 65A004.64Tip . . . . . . . . . . . . 65A003.16

Vertical tail (twin fin):Span, ft . . . . . . . . . . . . . . . . . . . . . . 0.71Area, ft . .. . . . . . .... . . .. 0.82Aspect ratio, per fin . 0 ....... ..... 2.45Taper ratio . . . . . . . . . . . . . . . . . . . . . . 0.358Leading edge sweep, ........ 460541Root chord, in . . . . . . . . . . . . . . . . . . 10.25Tip chord, in . . . . . . . . . .. . . . . . . . . 3.67Mean geometric chord, in .. . . .. . . . . . . . . 7.48Airfoil section .. . . . . . . . . . . . . . 65A004.5Toe-in angle, deg .. . . . . . .. . . . . . . . . 0.85Cant angle, deg outboard . . . . . . . . ....... . 5.0

Ventrals (twin ventrals, uncanted):Area, total, ft 2 . . . . . . . . . . . . . . . . . 0.13LE sweep, deg .......... . 40

10

NADC-81293-60

TABLE II - CONFIGURATIONS TESTED AND DATA FILE INDEX . CONFIGURATION e CONTROLS FILE

RIW22RV -111W22HVp+58 IIIW22HVp+10 HIV22HVp+15 DV22HVp-15

RIV22RV-JOr HIV22HVp+l5-JOr RIW22HV+7d-30r RIW22HVp+1S+7d-JOr ~W22HV+55ap+7d-30r HIW22HVD+l5+55ap+7d-30r

HIW22HV+55ap+12d-30r HBV22HVp+1S+55ap+12d-30r HIIV22HV+55ap+7d HIY22HVp+lS+55aD+7d

HIIV22HV-23h HBV22HVp+l5-23h 111V22HV-23h-30r HBV22HVp+15-23h-30r

HBW22HV-23h+55ap+7d-30r HBW22HVp+15-2Jh+S5ap+d-r HBW22HV-2Jh+55ap+7d HIW22HVp+l5-23h+55ap+7d HIW22HV-35hb HIW22HVD+15-3Shb

HBW22HV+eh HIW22HVp+l5+8h 111W22HV+8h+55ap+7d HIW22HVp+1S+8h+55aD+7d

HBW22HVCL£ANC HIV22HV-pocld HBV22HVp+l5-DOdd

HlcW22HV HlcW22HVp+5• HlcW22HVp+108 HkW22HVp+l5 lllcV22HVp-158

8 atatic (nb/2V•O) only

bte~ted for ~·20-90° only

-

6h 6 sp dea dea

0 0

I I , ..

+SS

I I I ' -23 ! 0 :

+ i

+55 I I ,. . ' -~s l i

+8 I i I I ' I I ~i~ 0 !

I 0

cRlove vane in, slats and flapA retracted

dchin pod r.aved

11

I

;

' I

I

6d I 6 8 NAME r dea dea de&

I 0 0 0 F14A

I +5 I F14AJ I +10 F14A4 I +15 I F14Al

I ' -15 i F14A2 I : I I i -30 ' 0 Fl4B4 I I +15 Fl485

+7 I 0 Fl4B14

~ i

j +15 F14B15 i I 0 F14B i ' +15 P14Bl

+t2 ! i 0 F14B16

i ' +15 F14Bl7 '- 0 0 Fl4B2 ' +IS Fl4B3 . I 0

! 0 F14B6

~ ! +15 Fl4B7 : -30 0 F14BC I

I I F14B9 I +15 +7 ! '

: 0 Fl4B22

I : +15 F14B23 0 I 0 F14Bl2

' i

I ! +15 F14Bl3 : 0 F14B28

' i ' +15 P14B29 0 ! 0 0 F14B24 I ' I +15 Fl4B25 '7 0 F14B26 +15 Fl4B27 0 0 Fl4CL

I 0 F14MP +15 Fl4MP1 0 F14C +5 F14C3

+10 F14C2 +15 F14C1

' -15 F14C4 edata were measured a t a cg location of 0.16c for all conf i gur ations

HADC-81293-60

~fl. CONFGURATIONTABLE II. - CONCLUDED FL

lb lop d 6r 8 l_ _ _on des don des __s

HcW22WV+55sp+7d-30r 0 +55 +7 -30 0 F14CSHlcW22UVp+5+55sp+7d-30r V +15 F14C6H~cW22HV-23h+sp+7d-r -30 F14C7H~cW22NV+15-23h+sp+-7d-r 1 +15 F14CO

UUBW50 0 0 0 0 0 F10D2MlW5Olp+15 II-0 +15 F10D3HBW5OV30r 300 F14D6HBV50RVp+lS-30r j I+15 F1071IMW509V+55sp+7d-30r +5 7 I0 MA4DEDB5VSOV 5+55sy 7d-30r -'.V. +15 710D1

HBW50HV+5sp+7d 0 0 F14D8UIVSOUHVp+l5+55sLP+7d - -- +15 F10D9

HBW50NV-23h -23 0 0 0 0 F104DIBW5ONVp+15-23h I -0 +15 F10D5HM15OV-23h-30r 300 F14DIONIVSOUM u15-23h-30r J1 F 14D1lHDV5OV-23b55p+7d-r +55 +7 0 71014A

* IW50Vp15-23h~sp+7d-r T +15 F10D15HNW08-2b+59+7 00 F10D12

HBW509VP+15-23b+55sp+7d V+15 F14D13

amnI vaacL~ 0 0 - 0- F10D16

HOV6SKV II0 F1412UIV68MVp+15 II1+15 F14E3HBV68NV+7d-30r I I -30 0 F14E

HINW68HVp+15+7d-30r V +5 lR

gloe anein sltsandflpsretracted *data were measured at a cg location

~gloe vne rtratedof 0.16c for all configurations

12

NADC-81293-60

TABLE III.- CONIFIGURATION DEFINITION

(1) (2~~ (3) 3h+5s(4) 3

(1) The leading H designates the NASA strain gauge balance used for the tests;

in this case the.HCFO3 balance.

(2) These characters define the configuration:

B body, including wing -glove, nacelles, ventral., and- chin pod.unless otherwise specified

c -canards; If the c Is omitted, the canard. are not present

W22- wing and A.LE, In degree.

H - horizontal tail

V - vertical tails

(3) The sideslip angle Is specified as p plus the angle In degrees. If nospecification is present, this Indicates zero sideslip angle.

* **(4) The control deflection. are specified as the deflection in degrees (signconvention as specified in the Symbols list), followed by letters designatingwhich surface. as follows:

h - symetrical horizontal tail deflection, 6 h

sp - spoiler deflection, 6 s

d - differential tail deflection, 6d

r - rudder deflection, 8r

If a surface is not designated, the deflection was zero. The maximum strtnglength for the configuration description is 25 characters. To meet thisrequirement for some configurations, the degrees of deflection for somecontrol surfaces were omitted; however, the sign of the deflection was re-tained. The deflections in such cases can be determined from Table 11.

13

MADC-81293-60

W- ! 00 14 -w 0

Q ..

0 0u0

N- *- -6 5% 5' °in N N C4 C4 N4

9d 00 #A .4G

k:.0"

15,+ C; C

Iad

4 0 14 4" N. N4 N Ne

i i_ _ _ _ _

++ll

0 C n 4 0% P%zAI+ P- +4~-

00 in in - 0

t4 0.

2 140 __ _ __ _ __ __ __

* A S,"11+ > " ' : 1 " + "4N 1+ 1 'm 2-' ++' + " i ': 1-;'2'22"m1 1+ 'i'i ""4 . " ""~"I+ i '" "" "+ +.. ".+"+i "i i

NADC-81293-60

U

AI en C4

zI>

00 0 0

(N S 0% rh-D. O

o > 0 0% m4-4_ _ _ _ _ _ __ _ _ _

00 in 0% en

> (n.-J ,.4 0 0% %0%

gel ,4 15

'4.4

I-44

LUw

ucn

N..;~~~-P C4 - l ' % 00

Ii4~Ij4J N ~15

NADC- 81293- 60

Figur e 1 . - Pho tograph of 1/12- scale model i ns talled on the rotary balance appa ratus .

16

NWD-81293.-60

A Slip ring housing8 Drive ShaftC Support borno Spin radius offset

potent iometerE CounterweightF Strut6 Angle of attack

pqsitiolifg motor

A

C

Spin axis4Velocity vector

(a) Side View Of Model-

Fiue2.- Sketch Of rotaybaneapat.

17

NADC-81293-60

A Slip rino housingB Spin radius offset

potenlti ometerC Lateral offset

drive gearsD Lateral offset

potent iomgeterE StrutF Sideslip angle

potentiometer-6 Sideslip angle

positioning motor

A

C

D

Spin axistVelocity vector

(b) Front view of model.

Figure 2.- Concluded.

18

a a--A

NADC-81293-60

64.1-

129

a. .. Sta.

6L. 9

NADC-81293-60

W.L. W.L. 14.0

~---__.-J

F.S. 7.7

F.S. 13.6

F.S. 20.2

Figure 4.- Ske~ch of canards tested on the 1/12-scale F-14 model.

20

D IS TR I B UTIO0N L IS T

REPORT NO. NADC-81293-60AIRTASK NO. AO3V-32D/OOlB/2F41-400-OOO

No. of Copies

Naval Air Systems Commnand .. .. ... .. ... ... .. ..... 4(2 for AIR-00D4)(2 for AIR-53011)

National Aeronautics And Space Administration, Dryden FlightResearch Center.......................2

(2 for M. Arebalo)....................Naval Air Development Center. .. . .. ... .. ... ......16

(I for 605)(I for 605DF)(2 for 60011)(3 for 6053DF)(3 for 8131)(6 for 6053:M. Stifel)

Defense Technical Information Center .. .. .. ... .. ..... 12Grummnan Aerospace Corporation, Bethpage. .. .. .. .. ...... 2

(2 for H. Beaufuere)

........ .... . .. ..... . ....