NASA Technical Memorandum 4782

F-15B/Flight Test Fixture II:

A Test Bed for Flight Research

David M. Richwine

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

CONTENTS

ABSTRACT .............................................................................. 1

NOMENCLATURE ........................................................................ 1

INTRODUCTION .......................................................................... 2

F-15B AERODYNAMIC TEST BED DESCRIPTION ............................................. 4

Vehicle Description and Flight Envelope ..................................................... 4

Flight Test Fixture II Description ........................................................... 5

Instrumentation ......................................................................... 5

Flight Test Conditions and Data Reduction ................................................... 8

RESULTS AND DISCUSSION .............................................................. 10

Flow Environment ..................................................................... 10

Surface Pressure Distributions ............................................................ 12

Boundary-Layer Measurements ........................................................... 20

Internal Environment ................................................................... 24

Bay Temperatures ................................................................... 24

Vibration Data ..................................................................... 25

Designing Fixture Experiments ........................................................... 28

Established Policies and Guidelines ..................................................... 28

Alternate Fixture Configurations ....................................................... 29

SUMMARY OF RESULTS ................................................................. 31

ACKNOWLEDGMENTS ................................................................... 31

REFERENCES ........................................................................... 32

APPENDIX A - STRUCTURAL LOAD CONSIDERATIONS FOR

ALTERNATE CONFIGURATIONS ....................................................... 33

APPENDIX B - AIRCRAFT STABILITY CONSIDERATIONS FOR

ALTERNATE CONFIGURATIONS ....................................................... 36

APPENDIX C - CHORDWISE PRESSURE DISTRIBUTION DATA ............................... 38

APPENDIX D - BOUNDARY-LAYER PROFILE DATA ......................................... 44

TABLES

C-1.

C-2.

C-3.

Steady-state chordwise pressure distribution data obtained in flight, Hp - 15,000 ft ................. 38

Steady-state chordwise pressure distribution data obtained in flight, Hp = 30,000 ft ................. 39

Steady-state chordwise pressure distribution data obtained in flight, Hp = 45,000 ft ................. 40

iii
C-4°

C-5.

C-6.

D-1.

D-2.

D-3.

D-4.

D-5°

D°6.

Chordwise pressure distribution data, obtained in flight, as a function of

angle of sideslip, Mach = 0.52 and 0.70 ................................................... 41

Chordwise pressure distribution data, obtained in flight, as a function of

angle of sideslip, Mach = 0.89 and 1.28 ................................................... 42

Chordwise pressure distribution data, obtained in flight, as a function of

angle of attack ....................................................................... 43

Steady-state boundary-layer profile data obtained in flight, Hp -- 15,000 ft ........................ 44

Steady-state boundary-layer profile data obtained in flight, Hp _ 30,000 ft ........................ 44

Steady-state boundary-layer profile data obtained in flight, Hp = 45,000 ft ........................ 45

Boundary-layer profile data, obtained in flight, as a function of sideslip,

Mach = 0.52 and 0.70 ................................................................. 45

Boundary-layer profile data, obtained in flight, as a function of angle of sideslip,

Mach = 0.89 and 1.28 ................................................................. 46

Boundary-layer profile data, obtained in flight, as a function of angle of attack

Mach = 0.54, Hp = 30,000 ft ............................................................ 46

FIGURES

1. Original k-TF installed on the lower fuselage of the F- 104 aircraft ................................ 3

2. FTF-II installed on the lower fuselage centerline pylon location of the F-15B aircraft ................ 3

3. F- 15B/FTF-II airspeed limitations ........................................................ 5

4. k-TF-II schematic diagrams .............................................................. 6

5. Removable airdata probe on the FTF-II .................................................... 8

6. Trim o_A and _ as a function of Mach number ................................................ 9

7. Wedge attached to side panel for shock wave/boundary-layer interaction experiment ................. 9

8. Flow visualization results obtained in flight, right side with tufts ................................ 10

9. Chordwise pressure distributions obtained in flight at specified Mach numbers

and altitude .......................................................................... 13

10. Chordwise pressure distributions obtained in flight at specified Mach numbers ..................... 14

11. Approximate regions of shock impingement on the surface of the FI'F-II

derived from flight data ................................................................ 15

12. Chordwise pressure distributions obtained in flight at specified FTF-II angles of attack;

Mach = 0.54, 13= 0 °, Hp = 30,000 ft ...................................................... 16

13. Chordwise pressure distributions obtained in flight at specified angles of sideslip;

Mach = 0.52, Hp = 15,000 ft ............................................................ 16

14. Chordwise pressure distributions obtained in flight at specified angles of sideslip;

Mach = 1.28, Hp = 45,000 ft ............................................................ 17

15. Chordwise pressure distributions obtained in flight at specified Mach numbers and

angles of sideslip, z/b' = 0.48 ............................................................ 18

iv
16.

17.

18.

19.

20.

21.

22.

23.

24.

25.

A-1.

A-2.

A-3.

B-1.

B-2.

In-flight relationship of M r with M._, outside the boundary layer, x/c = 0.90 ....................... 21

Boundary-layer profiles obtained in flight at specified Mach numbers and altitude .................. 21

Boundary-layer profiles obtained in flight at specified altitude and Mach numbers .................. 22

Boundary-layer profiles obtained in flight at specified FTF-II angles of attack;

Mach ---0.54, [3---0 °, Hp = 30,000 ft ...................................................... 23

Boundary-layer profiles obtained in flight at specified angles of sideslip .......................... 24

FTF-II bay temperatures obtained during a typical flight profile ................................ 25

FTF-II accelerometer data obtained in flight ................................................ 26

FI'F-II with aft fairing and flow visualization results obtained in flight, right side

with tufts, Mach = 0.70, Hp -- 45,000 ft .................................................... 29

F'IT-II configuration for the thermal protection systems durability program ....................... 30

F-15B on-board video camera locations, top view ........................................... 30

Load testing of the FTF-II mounted to the F-15B pylon ....................................... 33

The FTF-II reaction loads at the sway braces, flight- and ground-test data ........................ 34

Center of pressure locations determined from flight data as a function of Mach number .............. 35

The F- 15B dutch roll mode damping obtained from NASA Dryden batch simulation

(CAS off) with and without FTF-II; Hp = 30,000 feet, aircraft weight of 36,200 Ibm ................ 37

Comparison of predicted and flight test values for F-15B dutch roll mode damping,

CAS off, FIT-II on for flight, FrF-II on for simulation ....................................... 37

ABSTRACT

NASADrydenFlightResearchCenterhasdevelopedasecond-generation flighttestfixtureforuseasageneric

testbedfor aerodynamicandfluidmechanicsresearch.TheFlightTestFixtureII (VIT-II) is a low-aspect-ratio

verticalfin-likeshapethatismountedonthecentedineoftheF-15Blowerfuselage.Thefixtureisdesignedforflight

researchatMachnumberstoamaximumof 2.0.TheVIT-II is a composite structure with a modular configuration

and removable components for functional flexibility. This report documents the flow environment of the fixture, such

as surface pressure distributions and boundary-layer profiles, throughout a matrix of conditions within the

F-15B/VYF-II flight envelope. Environmental conditions within the fixture are presented to assist in the design and

testing of future avionics and instrumentation. The intent of this document is to serve as a user's guide and assist in

the development of future flight experiments that use the FTF-II as a test bed. Additional information enclosed in

the appendices has been included to assist with more detailed analyses, if required.

NOMENCLATURE

A S

b'

C

Cn_

Cp

Cpmax

Cp*

CAS

F

Y

FrF

FTF-II

g

Hp

KCAS

l

M

M r

M_

ORD

Pw

P_

FTF-II side area, 3424 in 2

span height from the top of the FTF-II, 32.0 in.

chord length from the leading edge of the FTF-II, 107.0 in.

static directional stability, deg -1

pressure coefficient, (p - p._)/q._

maximum pressure coefficient

critical pressure coefficient

control augmentation system

sway pad reaction load, lbf

side force, lbf

Flight Test Fixture (F-104)

Flight Test Fixture II (F-15B)

acceleration of gravity

aircraft indicated altitude, ft

knots calibrated airspeed

reference length with turbulent boundary layer, x - Ax, in.

aircraft indicated Mach number

maximum local Mach number at rake, x/c = 0.90

moment on FTF-II caused by side loads, in-lbf

Fq_F-II noseboom free-stream Mach number

Objectives and Research Document

reference static pressure for rake, lbf/ft 2

FTF-II noseboom free-stream static pressure, lbf/ft 2
PCM

PSD

q**

Re

nTIS

T

rl

u

Ur

w

x

x/c

Y

Z

z/b '

0_ n

pulse code modulation

power spectral density

FTF-II noseboom free-stream dynamic pressure, lbf/ft 2

Reynolds number

root mean square

absolute temperature, R

FTF-II total airstream temperature, x/c = 0.96, R

local wall temperature for rake, x/c = 0.90, R

axial velocity component within the boundary layer, ft/sec

maximum local velocity for rake, ft/sec

width, FTF-II, 8.0 in.

distance from leading edge, in.

percent chord, FTF-II

distance from wall, in.

distance from leading edge, in.

percent span, FFF-II

FTF-II noseboom local angle of attack, deg

aircraft indicated angle of attack, deg

FTF-II noseboom local angle of sideslip, deg

boundary-layer thickness, in.

approximate distance from origin of turbulent flow to nose, 12.84 in. (x/c. = 0.12)

damping ratio

undamped natural frequency, rad/sec

INTRODUCTION

In 1963, NASA Dryden Flight Research Center (NASA Dryden) constructed a low-aspect ratio fin-like shape

called the Flight Test Fixture (FFF) for installation on the lower fuselage of an F- 104 aircraft (fig. 1). Since that time,

the FIT has been used as a generic test bed for more than 20 aerodynamic and fluid mechanics research experiments.

The capabilities and utility of the FFF have been documented frequently. 1-8 In recent years, NASA Dryden has

phased out the aging F-104 aircraft. An effort has subsequently been made to develop a larger and more versatile

flight test fixture and identify a new test bed aircraft with expanded capabilities relative to the F-104 FTF. An F-15B

airplane was chosen as the test bed aircraft for the new flight test fixture because of its availability at NASA Dryden

and potential for good flow quality because of the relatively smooth lower fuselage surface. In addition, the

ability of the F-15B airplane to reach Mach 2.0 and its in-flight refueling capability were considered to be very

desirable features.

The new flight test fixture was installed on the lower fuselage centerline pylon location of an F-15B aircraft and

is called the Flight Test Fixture II (FrF-II) (fig. 2). Envelope expansion of the F-15B/FFF-II configuration was

2
Figure1.OriginalFIT installedonthelowerfuselageof theF-104aircraft.

Figure2.TheFTF-II installedonthelowerfuselagecenterlinepylonlocationof theF-15Baircraft.

completedinathree-phaseeffort.Phaseoneconsistedof flyingtheFTF-IIuninstrumentedontheF-15Bairplane

toperformenvelope-expansion andqualitativeflow-qualitystudies9toamaximumofMach1.3.Directionalstability

wasevaluatedby performingrudderdoublets.Flowqualitywasdocumentedusingtuftsto definesurfaceflow

streamlines.

Phasestwo(Mach< 1.8) and three (Mach > 1.8) were performed with an instrumented FFF-II that included

airdata and total temperature probes. The airspeed envelope was expanded to the current aircraft performance

and FFF-II design limits of 600 knots calibrated airspeed (KCAS) and Mach 2.0. Envelope expansion continued,
includinginvestigationsrelatedto directional stability, 10 structures, and aerodynamics. Calibrated load cells

were located at the sway brace pads, and preliminary surface pressures were used to monitor in-flight lateral loads

and the distribution of these loads. Final envelope expansion of the F-15B/FTF-II configuration was completed in

December 1994.

Aerodynamic and environmental characterization of the F-15B/FTF-II began in April 1995. The primary

purpose of this report is to document the flow environment of the fixture, such as surface pressure distributions and

boundary-layer measurements, throughout a matrix of conditions within the F-15B/FTF-II flight envelope. Environmental conditions within the fixture are presented to assist in the design and testing of future avionics and instrumentation within the fixture. Several guidelines and recommendations are presented to assist in the development of

future experiments that may utilize the FTF-II as a test bed.

F-15B AERODYNAMIC TEST BED DESCRIPTION

The NASA F-15B aerodynamic test bed is the combination of the F-15B with the FTF-II installed and is

referred to as the F-15B/FTF-II. The following section provides a detailed description of the F-15B/FTF-II configuration, including a description of the F-15B aircraft and capabilities and the FTF-II hardware and instrumentation,

and an overview of the flight experiment.

Vehicle Description and Flight Envelope

The NASA F-15B airplane is a two-seat version of the F-15 aircraft, which is a high-performance, supersonic,

all-weather, air-superiority fighter built by McDonnell Douglas Aerospace (St. Louis, Missouri).ll The aircraft is

powered by two Pratt & Whitney (West Palm Beach, Florida) F100-PW-100 turbofan engines with afterburners. The

F-15B airplane is 63.7 ft long with a 42.8-ft wingspan and a basic operating weight of 27,500 Ibm. The aircraft appearance is characterized by a high-mounted swept-back wing, twin vertical stabilizers, and an elevated cockpit to

enhance visibility. Primary flight control surfaces consist of conventional ailerons, twin rudders, and stabilators.

These surfaces are controlled by a hydromechanical system and an electrical control system called the control

augmentation system (CAS).

The F-15B/FTF-II loading is similar to the F-15 Air Superiority "A" loading,Â 1 with the FIT-II at the centerline

tank location. The FI'F-II is installed on the F-15B centerline pylon in a manner similar to operational centerline

tanks. The maximum projected weight of the FTF-II, including all research systems, is 500 Ibm. The FI'F-II is

smaller, weighs less, and has a lower aerodynamic drag than an empty 610-gal centerline tank that is routinely carried by the F-15 aircraft. As a result, issues such as aircraft stability, performance, and loads were addressed by comparison to the Air Superiority "A" loading for the F-15 aircraft. The FTF-II was designed for flight testing within

the operational F-15 airspeed envelope (fig. 3) with an additional limit of Mach 2.0 because of aerodynamic heating

considerations on the VIT-II composite structure. Flight testing of the F-15B/F'IT-II configuration with the current

F100-PW-1 O0 engines shows that steady-state conditions can be achieved at altitudes to a maximum 60,000 ft and

airspeeds to a maximum 600 KCAS and Mach 2.0.

If airspeeds to the maximum aircraft limits of 800 KCAS or Mach 2.3 are desired, further evaluation of several

issues would be required. Overall performance of the F-15B airplane would need to be improved with a minimum

upgrade to the Pratt & Whitney F100-PW-220 engines. Increased Mach numbers would require the installation of a

VI'F-II nose section designed for higher levels of aerodynamic heating with an aerodynamic shape more appropriate

for improved flow quality at high speeds. Some additional analysis and verification of FTF-II loads and aircraft

stability would also be required.
60 x 103

5O

(aircraft

limit)

40

Altitude,

ft 30

20

10

ml

01 )0 200 300 400 500 600 700 800 900

Calibrated airspeed, knots

Figure 3. F-15B/FTF-II airspeed limitations.

960729

Flight Test Fixture II Description

The FTF-II is a low-aspect-ratio fin-like shape (fig. 4) 107.0 in. long, 32.0 in. high, and 8.0 in. wide that has an

elliptical nose section and blunt trailing base. The FTF-II is a composite modular structure and consists of the basic

avionics pylon with a replaceable nose section, side panels, and vertical test article that allow the configuration of

the FTF-II to be modified to satisfy a variety of flight test requirements. The upper 19 in. of the FTF-II, the avionics

pylon, is a permanent structure that houses avionics, research instrumentation systems, and other support equipment

common to most flight experiments. The lower 13 in. of the FTF-II is the vertical test article that, in the current configuration, matches the contour of the upper avionics pylon. The vertical test article is removable and may be replaced by other aerodynamic shapes. Normally, only instrumentation specific to individual research experiments is

installed in the lower vertical test article.

All FTF-II side panels are removable with quick internal access through the four left side panels that use external

fasteners flush to the surface. The two right side panels extend the length of the fixture and are attached using internal

fasteners to minimize discontinuities for aerodynamic experiments such as boundary-layer experiments or surfaceflow visualization studies. The elliptical nose section is located in the forward 18.8 in. (one-half major axis) and can

be removed independently from either the avionics pylon or the vertical test article. A removable pylon interface fairing and a splitter plate have been installed onto the upper portion of the fixture for improved flow quality. Provisions

were also made for a removable 7-in. by 10-in. test panel on the right side of the vertical test article nose section.

This test panel has already been used for instrumentation studies that require laminar flow or aerodynamic heating.

A removable airdata probe is located at the bottom of the vertical test article nose section.

Instrumentation

The FTF-II data acquisition capability is primarily located in the fixture avionics pylon and uses a 12-bit pulse

code modulation system capable of multiplexing data at sample rates as many as 200 samples/sec, depending on the

number of data channels. Flight data are transmitted to the ground for storage and postprocessing. A tape recorder
locatedintheF-15Bairplaneisusedasabackupcapabilityortorecordhighfrequencydataif requiredforaspecific

flightexperiment.

Fourloadcells(fig. 4(b))werelocatedat theswaybracesto helpmonitorFrF-II installation preloads and

in-flight side loads. Each of the load cells was rated at a maximum of 10,000 lbf compression and oriented to isolate

lateral reaction loads as much as possible. Ground calibration load tests were conducted on the F-15B pylon/

Splitte_lonlcs

Ilipticel no_

section

Vertical test article

(a) Exploded isometric view.

Left side panels

Bey Dimension (In.)

A1 12.0 x 7.0 x 7.0

A2 12.0 x 18.7 x 7.0

A3 12.0 x 20.8 x 7.0

A4 12.0x23.6x7.0

T1 8.5xT.0x7.0

T2 8.5x18.7x7.0

T3 8.5x20.8x7.0

T4 8.5x23.6x7.0

Pdme_ utili_

Instrumentation

Power, PCM

Instrumentation

Experiment

Alrdeta

Experiment

Experiment

Experiment

960730

Sway braces/

load cells (4) -_

BE "BB

O

im mm

Suspenslon lugs (2)

O

Top view

\

"_-Avionlcs access

•,,----18.8 -_

,,m

35.7

C : 107.0

Left side view

52.5

(b) Three view.

Figure 4. FTF-II schematic diagrams.

m

Aft view

b':'32.0

960731
FTF-II installation to correlate known side loads on the fixture to in-flight loads at the load cells. Appendix A

provides details of this ground calibration.

Flush static-pressure orifices (0.030 in. diameter) are located at three spanwise (z/b') locations and distributed

chordwise (x/c) (fig. 4(b)). The upper leading edge of the FTF-II was used as the reference z = 0 (+ down) and

x = 0 (+ aft) locations, respectively. The static-pressure orifices are located on both sides of the nose section forward

of x/c = 0.15 and on only the left side aft of this chord location. A boundary-layer rake located on the left side of the

FTF-II (figs. 4(b) and 4(c)) was used to measure boundary-layer profiles at z/b' = 0.70 (probe near wall) and

x/c = 0.90. This rake was canted to accommodate a fine probe resolution near the wall surface (fig. 4(c)) with the

(c) Detail of boundary-layer rake.

Figure 4. Concluded.

outermost probe located 2.5 in. from the surface. Appendix D includes specific probe locations with the tabulated

rake data. Pressure measurements were obtained using four temperature-controlled electronic scanning pressure

transducers at a sample rate of 6 samples/sec. Surface and rake pressures were measured with +720 lbf/ft 2

and +2160 lbf/ft 2 differential transducers, respectively, referenced to the noseboom static pressure with estimated

precision within +3 lbf/ft 2.

Figure 5 shows a removable airdata probe that provides static and total pressure to measure local flow conditions

just forward of the fixture. The airdata probe also measures FTF-II noseboom local angle of attack, o_, and local

angle of sideslip, 9. Total temperature is obtained from a probe located on the aft left side of the fixture (fig. 4(b)).

Documentation of acceleration data and internal bay temperatures was conducted to assist with environmental

design guidelines for future flight experiments. Accelerometers were located in the bottom of the FTF-II vertical

test article midchord to document local three-axis accelerations during various flight conditions. Internal bay temperatures were also obtained to document typical bay temperatures over a variety of ground and flight conditions.

7
Figure5.Removableairdataprobeon the FTF-II.

Outside wall temperature was measured near the base of the rake at x/c -- 0.90 (fig. 4(c)) to assist with boundarylayer calculations. All of these parameters, including an additional canted boundary-layer rake for the right side of

the fixture, are available for future flight experiments.

Flight Test Conditions and Data Reduction

Flow quality data were primarly taken at 1-g flight conditions and a matrix of nominal altitudes of 15,000 ft,

30,000 ft, and 45,000 ft in increments of approximately Mach 0.1 throughout the airspeed envelope of the aircraft.

Maneuvers such as level accelerations and decelerations and sideslips were conducted in a manner to minimize aircraft fuel consumption while maintaining proper position within the supersonic corridor. For example, at an aircraft

indicated altitude, Hp, of approximately 45,000 ft, a level acceleration was first performed to a maximum Mach 1.5,

followed by a level deceleration and sideslips at specified Mach numbers. Data at greater than Mach 1.5 were obtained by diving from a higher altitude to Mach 2.0 in the most efficient manner possible and then performing a level

deceleration at an altitude of 45,000 ft.

Other parameters such as aircraft indicated angle of attack, t_A, and angle of sideslip tend to be limited by aircraft

performance. Figure 6 shows that trim c_A slowly decreases and approaches values between 2 ° and 0 ° with increasing

Mach number for altitudes between 30,000 and 45,000 ft. Further discussion on the FTF-II noseboom angle of attack will be included in the "Flow Environment" section of this report. At subsonic speeds, local angles of sideslip

as high as +8 ° were attained during most maneuvers except at dynamic pressures greater than 500 lbf/ft 2. The F-I 5B

flight controls limit the rudder deflections at supersonic speeds. Maximum local angles of sideslip as high as +4 ° are

attainable at approximately Mach 1.3, but flight data indicate that maximum local angles of sideslip are reduced to

approximately +2 ° at higher Mach numbers.

Future flight experiments may require increased angles of incidence to meet specific research objectives. These

increased angles can be achieved through the use of pitchups or sideslips depending on the experiment location and
10

8

6

4

2

Angle

of attack, 0

deg

-2

-4

-6

-8

- 10.4

i I I I I '

i

tt ,' p_ _' _p CxA , 30k ft

...... .4...... _ ..... .4...... _ ..... .4-- - --- c_-, 45k ..ft --

\ ' 1 ', 1 i

A

' \ _ (_, 15k ft • I I I I

.... -__-L .... __L- ..... .L ...... l ..... J.__

, t i + _ 5,30kft -- \

•, _ , a , _ (x, 45k ft

...... i_ .... r [-- --f ...... r ..... T ...... _...... T ......

I • i %. I I i I I

: \', -. ,, ,, : : :

+L i.+L i G. _ " _i .11" + i

_.=, _b..,==, __,__ _._______,__ ..... ___-.____ ..... .4...... L-..... • ......

\ \I I +411,+ I+,/\I I I I I

........ ____1 t _ I!...... !t ......

I I r I I I

I + _ I I I

...... "1 ...... r- ..... "1 ...... 4- ..... "1- --r- ..... -r ......

I I _ I I I P

I I t I I I I

...... -4 ...... _- ..... 4 ...... _- ..... 4 ......... 4- ......

I I 1 I I

I I I 1 I

...... .J ...... (- ..... -L ............ J. ...... L ..... L___

I 1 I r I I

: I ....

• 6 .8 1.0 1.2 1.4 1.6 1.8 2.0

Mach number

Figure 6. Trim O_A and (_ as a function of Mach number.

960734

configuration of the FIT-II. Typically, these increased angles of incidence can only be reached at subsonic speeds

because of flight control limitations for the F-15B aircraft 10,11 that limit control inputs at supersonic speeds. At

Mach 0.5 and an altitude of 30,000 ft, the aircraft can momentarily achieve indicated angles of attack of approximately 20 ° without significant changes in other flight conditions such as altitude or Mach number. This limitation

results in a change in local flow angle on the fixture noseboom of-1.9 ° < et < 4.0 ° and indicates that the flow around

the FTF-II is fairly insensitive to aircraft angle of attack. As a result, obtaining local angles of incidence in the pitch

axis of the aircraft that are greater than approximately 6° can only be attained by mounting a horizontal surface to

the fixture side panel, such as the wedge shown (fig. 7). Larger angles of incidence in the pitch axis can also be

Pylon Interhice Miring

_0ra_

Figure 7. Wedge attached to side panel for shock wave/boundary-layer interaction experiment.

9
achievedby mountinganobjectto bottomof theavionicspylonatapredeterminedangleof attack.TheFIT-II

combinedwithsideslipsatsubsonicspeedscanbeusedtoobtainarangeofanglesofincidence+8 ° in the yaw axis.

The data from all sensors were averaged to minimize errors caused by slight fluctuations in conditions and signal

noise. Data from relatively steady maneuvers such as level accelerations and decelerations were averaged over a

1.0-sec interval. More transient maneuvers such as sideslips and pitchups were averaged over a 0.4-sec interval

because of rapid changes in flight conditions that would have otherwise been lost in the data reduction process.

RESULTS AND DISCUSSION

Flow conditions at the centerline tank location under the F-15B airplane were characterized by a variety of

measurements on the FTF-II. FTF-II noseboom airdata were combined with surface pressure and boundary-layer

measurements to survey aerodynamic flow quality. Environmental conditions within the FTF-II were documented

as a baseline to assist in the design and testing of avionics and instrumentation for future experiments.

Flow Environment

Flow visualization studies on the FTF-II (fig. 8) indicate that relatively chordwise flow conditions exist over the

FTF-II. Flow visualization data are substantiated by measurements of the o_(fig. 6). At subsonic speeds (fig. 6), the

o_ranges from 1° to -3 °, depending on the specific flight conditions. At supersonic speeds, downwash over the fixture

increases with increasing Mach number to the point where a is approximately -9 ° at Mach 2.0 (fig. 6). Flow

visualization studies (fig. 8) tend to support these results for subsonic and supersonic speeds.

(a) Mach = 0.70, Hp ---45,000 ft, [3= 0 °.

i 960736 i

EC96 43398-25

(b) Mach = 1.00, Hp = 45,000 ft, [3= 0°.

Figure 8. Flow visualization results obtained in flight, right side with tufts.

10
EC9643398-27

(c)Mach= 1.10,Hp = 45,000 ft, _ - 0 °.

(d) Mach = 1.20, Hp = 45,000 ft, _ = 0°.

EC96 43398-29

Indication of shock

impingement on surface

(e) Mach = 1.25, Hp -- 45,000 ft, _ -- 0 °.

960737

(f) Mach = 1.30, Hp -- 45,000 ft, 13= 0 °.

Figure 8. Concluded.

EC96 43398-31

11
Althoughtheflow over the FTF-II side panels is generally chordwise, subsequent pressure-distribution data and

boundary-layer profiles reveal the effects of shocks from the vicinity of the leading edge of the centerline pylon and

fuselage junction. This effect is observed in the discontinuity of the tufts (figs. 8(c) to 8(e)) with the most apparent

effects at Mach 1.25. At Mach 1.30 (fig. 8(f)), more chordwise flow conditions reappear. Shocks emanating from the

engine inlets were also believed to be the cause for discontinuities in the airdata between Mach 1.0 and 1.4. These

anomalies will be highlighted in later sections of this report.

The current F- 15B/VYF-II airspeed envelope for airspeeds to a maximum of 600 KCAS and Mach 2.0 results in

a maximum range in unit Reynolds numbers from 1.0 x 10rift to 6.0 x 106/ft. As previously reported, 12 the lower

limit for the critical Reynolds number for boundary-layer transition for a flat plate at low speeds is approximately

3.2 x 105, although values greater than 1.0 x 106 can occur for very disturbance-free flow. Thus, considering the unit

Reynolds number range for the fixture, turbulent flow can be expected to begin naturally within the first foot

(x/c < 0.11) of the FrF-II nose section at low-speed conditions. Previous FTF-II flight test results 9 using liquid

crystals found natural transition locations on the nose section between x/c = 0.10 and x/c = 0.12, and as far aft as

x/c = 0.16 at Mach 1.1. Forced transition using artificial surface roughness should be used ifa fixed transition point

forward ofx/c = 0.10 is desired for a specific experiment.

Surface Pressure Distributions

Pressure coefficients, Cp, were obtained using the FIT-II free-stream noseboom static pressure, p_, and

dynamic pressure, q**, for reference conditions. Appendix C documents chordwise pressure distributions obtained

on the left side of the fixture at z/b' = 0.16 (upper), z/b' = 0.48 (mid), and z/b' = 0.86 (lower) for the entire matrix of

flight conditions. The maximum value of the pressure coefficient, Cpm_, at the stagnation point for subsonic flOW 13

and behind a normal shock wave for supersonic flow 14 was calculated and included for steady-state maneuvers at

an altitude of 45,000 ft for comparison and validation of the FTF-II nose section pressures at x/c = 0.0 (table C-3).

Figure 9 shows typical FTF-II chordwise pressure distributions for Mach 0.7 and Mach 1.5. Pressure coefficient

values have been offset -0.4 and -0.8 for spanwise locations of z/b' = 0.48 and z/b' = 0.16, respectively, on the chordwise pressure distribution plots for comparison. As expected, no significant Reynolds number effect at the selected

altitudes was observed. Obvious features from the chordwise pressure distributions include the suction peak from

the contoured nose section, slight base area effects for x/c > 0.96 (fig. 9(a)), and the presence of a weak shock from

the centerline pylon at x/c -- 0.18 (fig. 9(b)).

Figure 10 shows differences in the chordwise pressure distributions as a function of Mach number. In the subsonic region (fig. 10(a)), chordwise pressure gradients are small between 0.20 < x/c < 0.90. At transonic speeds

(figs. 10(b) to 10(d)), chordwise pressure distributions increasingly show the effects of compressibility, and the formation of the normal shock from the FTF-II surface is observed at Mach 0.89. Figures 10(b) and 10(c) show the

value of the critical pressure coefficient, Cp*, which corresponds to sonic flow. Figure 11 shows the approximate

regions where shock impingement locations were identified on the surface of FTF-II. The normal shock from the

FTF-II surface at Mach 0.89 is located slightly upstream ofx/c = 0.21. From Mach 1.00 to 1.39, the effects of the

weak shock from the centerline pylon are seen at the three spanwise locations. At Mach > 1.39 (figs. 10 and 11),

effects of the shock move aft and weaken. A more detailed analysis of this region of weak shock interactions would

require additional instrumentation.

Smaller chordwise pressure gradients are again present for Mach > 1.60 (figs. 10(f) to 10(h)). The slight rise in

the surface pressure at z/b' = 0.86 and x/c = 0.90 at supersonic speeds is believed to be caused by the presence of the

boundary-layer rake near this location and the resulting rise in local static pressure. The influence of the rake on local

flow is typical of a local disturbance that may occur if external hardware is mounted on a side panel of the FTF-II

for future experiments.

12
Cp

- 1.0

-- .8

-- .6

-.4

--°2

0

-- .2

0

I I i

I I

, , _ np= IOKn

I I

.......... +.................... _---- + Hp=30kft ---

I

......... ctio_n_p_e_ak_............ _ Hp = 48k ft 5; ...............

Upper

z/b' = 0.18

Mid

z/b' = 0.48

-.2

0

.2

Lower

z/b' = 0.86

0 .2 .4 .6 .8

xlc

1.0

960738

(a) Mach= 0.7.

- 1.o i ', '

_ _ Hp = 15k ft

- .8 .......... + ......... -_.......... F--- ....-.6--- Hp- 30k ft---

t

' _ _ Hp = 45k ft

l I

- .6 .......... f .........................................,

r :

.......... T .......... I.................... "l ..........

:/-- Weak shock _ _,

--

0 .2 .4 .6 .8

x/c

Upper

z/b' = 0.16

Mid zJb = 0.48

Lower z/b = 0.86

1.0

960739

(b) Mach = 1.5.

Figure 9. Chordwise pressure distributions obtained in flight at specified Mach numbers and altitude.

13
Cp

- 1.0

--.8

--.6

--.4

--.2

0

-.2

0

-- .2

0

.2

I I I _ 1 )

.... ---------------.---.--- -_-z,,_' = 0.48 -

I I t I I I n

_ t , _ _ --,e.--Z_ = 0.86 I

.... "----'-----'___,___,___,___,___,___,___

I I t I t F I i i

I l I l I I I I I

---,---,---,---,---,---,---,---,-,-

I I I I I 1 i b I

I I I I I I I I

.-----.-----.------.-----.----.-----.---.----

I I I I I i I I

_,,: :_; _l d, _l

I I I I I I I I I

I I I I I I r I I

_ _ _I___ I___ I___.J.__ _.L __ _ L_ __ L__ _ L__

I I I I I I I

J T i i

I I

0 .2 .4 .6 .8 1.0

x/c

960740

(a) Mach = 0.39, Hp -- 15,000 ft.

Cp

- 1.0

--.6

--.6

--.4

-- .2

0

-.2

0

--°2

0

.2

I i 1 i T _ , , , r

: : l : * ' -----z_l:)' = 0.16 I

I I

.... "--'---'---'---'---'-----'' = 0.48 -

i I i I I t I

' , , , _ , --e--Z/b' = 0.86 I

.... .--.___.___.___.___.___

I I I I I I I I i

I I I I I I I I I

.... "-----,___,___,___,___,___,_ _

I I 1 I I I

/ , "_; _"}_Cn*ir 0.43 ; ; ;

I IV I I I I I

L..... __'____.' ........... ___

I I I I I I I I I

.__,___,_,___,___,___,___,___,___,_,_

I I I l- " I " .

I I I I I I I F I

I I I I I I I I I

0 .2 .4 .6 .6 1.0

x/c

960741

(b) Mach = 0.80, Hp= 15,000ft.

Cp

- 1.0

--°8

--.6

--.4

--.2

0

-.2

0

--.2

0

.2

i i i i i ! .:.-,-zro' ,1).161

....._---"...._....i....!....!---.-.-,,o..o.48 1

....."___- ....;....i....i....;...._':.°:__ 1

.._____?::: .... -_.... ; .... i:" _.... ; .... ;_:.... :.....

,

0 .2 .4 .6 .8 1.0

x/c

960742

(c) Mach = 0.89, Hp = 30,000 ft.

0 .2 .4 .6 .8 1.0

x/c

960743

(d) Mach = l.O0, Hp= 30,000 ft.

- 1.0 .........

! I --,O,--- '-- , '_ ', ', , , z/b - 0.16

- .6 I----_'-- -_----_----, ........... -e-zro' = 0 48 -

i i I I P P

/ ...... ---,-,._'=o._ I

- 6 '---'--'---'---'---'---'---'-----------'

- I4 ___l___j ....... .............. _-___L___L___L___ I

F I ; d

- .2

Cp 0

-.2

0

P f _ t t I

iI _ uI iI F I I I

-.2,_ -'_-"_--'-------_---,---L---,---_-_-

I I I I t I _ i I

0 .2 .4 .6 .8 1.0

xlc

960744

- 1.0

--.6

--.6

--.4

.2

Cp 0

-.2

0

--.2

0

.2

0 .2 .4 .6 .8 1.0

xlc

960745

(e) Mach = 1.39, Hp = 30,000 fl. (f) Mach = 1.60, Hp = 45,000 ft.

Figure 10. Chordwise pressure distributions obtained in flight at specified Mach numbers.

14
- 1.0

--.8

--.6

--.4

--.2

Cp 0

_.2

0

--.2

0

.2

i ! i i ! ;; _-,_,=o.18[

.... _____1.... , .... _.... •.... ---- --=--z_' = 0.48 --I

i i t i t

...... --,e--z_ - 0.86 |

i = i _ t b * i *

.....

i

m---m

,

.... m

b

.... _

=

.... •

i

....

_

r ....

i

r ....

i

m ....

i

_--- 1 q , n , n i * t

• I i n l n I ; I

i t = t = i t m i

/

.... I L I I l v

0 .2 .4 .6 .8 1.0

xlc

960746

(g) Mach = 1.80, np -- 45,000 ft.

Figure 10. Concluded.

- 1.0

-- .R

--.4

Cp 0

-- .2

0

--.2

0

! i i i i :; -.-z_.=o.161

..... "---4 .... _'.... " .... ; .... r--- _z/b' ffi 0.48--[

: , : : : --.-z_,= o._ |

..... r- ----_ .... _ .... _ .... v .... r .... _ .... r .... _....

..... ;---_ .... _.... _.... T..... - ....... r .... '....

.................. _ .... _ ........................

!---- n ,_ --

, ', -_" ', ',---,- i , , --,,-

,

0 .2 .4 .6 .8 1.0

xlc

960747

(h) Mach = 2.00, Hp - 45,000 ft.

Aircraft fuselage

=/

z/b' = 0.16

z/b' = 0.48

zlb' = 0.86

Centerllne pylon

Left aide view of FTF-II

weak

ahockregion

I I I I I I I I I I I

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

xlc

960748

Figure 11. Approximate regions of shock impingement on the surface of the FTF-II derived from flight data.

Figure 12 shows chordwise pressure distributions at approximately Mach 0.54 are essentially insensitive to

changes in ct from -1.85 ° to 4.05 °, as expected. Figures 13 and 14 show differences in the chordwise pressure distributions as a function of sideslip. At Mach 0.52 (fig. 13), the most significant feature of the chordwise pressure

distributions is the suction peak followed by recompression on the leeward side of the fixture. As 13approaches -8.0 °

(windward side), pressure coefficient values approach zero for x/c > 0.20. Figure 14 shows pressure distributions for

Mach 1.28 as 13varies approximately +3.6 °. Spanwise effects as a function of 13appear to be minor.

Because the differences in the chordwise pressure distributions for the three spanwise locations are relatively

minor, the spanwise location of z/b' = 0.48 was selected to present chordwise pressure distributions as a function of

Mach number and sideslip. Pressure coefficient values (figs. 15(a) to 15(c)) have been offset -0.8, -0.6, -0.4,

and -0.2 for 13of 8.0 °, 4.0 °, 0 °, and -4.0 °, respectively. Figure 15(d) shows pressure coefficient values that have been

15
Cp

- 1.0

-.8

.6

-.4

--.2

0

-.2

0

-.2

iit

, , , ---I-- et - 1.85 °

............. _ ............ _............ _- -

a i i

: : : + (_ 0.30 °

............. _-............ _-............ 1___ -"'0"-- (_ 2.19 °

_. 4.05 °

t t i

t t p

............. T- ............ . ............ r- ............ _ ...........

f _ t

f

0

: a i

"20 .2 .4 .6 .8 1.0

xlc

960749

Figure 12. Chordwise pressure distributions obtained in flight at specified FTF-II angles of attack; Mach = 0.54,

f5= 0% Hp = 30,000 ft.

Cp

- 1.0

--.8

--o6

--.4

--._)

0

0 I

- .2 I

0

i _ i ; /

........ _......... i ......... ;- .-*-- z/b' = 0.16 - 4

/- Suction peak : i + z./b' - 0 48 1

__ '-_ ......... !......... _--.- _'= o._ -1

--'_-'_7- Recompmssion :-......... :-........ ]

/

! ! ! /

.2 .4 .6 .8 1.0

xlc

Ca) 13 = 8.04 °, leeward.

960750

Figure 13. Chordwise pressure distributions obtained in flight at specified angles of sideslip; Mach = 0.52,

Hp= 15,000 ft.

16
- 1.0

--.8

1 °6

1.4

--o2

Cp 0

.2

0

.2

0

.2 0

! i i _ ii

......... --,......... -,......... ,-..---- z/b' = 0.16

' ' ' _ z/b'=0.48

, , , --4- z/b'=0.86

t i

.2 .4 .6 .8 1.0

x/c

(b) 13= 3.95 °, leeward.

960751

- 1.0

-- °6

--.6

I °4

Cp

-.2

0

--°2

0

0

.2 0

............................. i----=-- z/b'= 0.16-

......... _;.......... ......... i_._ z/b' = 0.48. J

! -.-_,_-o,_t

q _ i j

J

" ', i :

.2 .4 .6 .8 1.0

(c) 13= o.oo °.

X/C

960752

Cp

- 1.0 ; ; ; ;

J t t

.8 ................. ; ......... '-" z/b' : 0.16

t t

: I : _ z/b'=0.48

, _ z/b'=O.86

J

- .4 ........ " ......... " ......... " ......... _-.......

1.2 - ....... ; ......... " ......... :-......... " .......

o_ :_

-" _i-: :! _ i-_

O i 1 .

-= _:_i:::i :...._

6 i . . . .

i i i

t t i

.2

0 .2 .4 .6 .8 1.0

x/c

960753

(d) 13=-3.98 °, windward.

-.2

0

-- °2

0

.2

0

Figure 13. Concluded.

-1.o| i ! i i,

-.81- ........ "......... " ......... "-- --e-- Z/b = 0.16 --

I, ! ! ": -.- zr,'=o._

-.6[ ........ _......... _......... 7-'- zr,'=o.M-

-.,[........ _....... _......... i....... _.......

-=_ ...._....?:j

Cp 0

i i i

i i i _1

.2 .4 .6 .8 1.0

x/c

960754

(e) 13= -8.00 °, windward.

- 1.0

--°6

-- °4

Cp 6

-.2

0

-- .2

0

.2

i i i'I

......... _......... i ......... _--.- z_ =o.16

,, : , --_- z/b' = 0.48

......... _ ......... _ ......... "- -4.-- z/b' = 0.86

_ ', ! i :

0 .2 .4 .6 .8 1.0

x/c

960755

(a) 13= 3.57 °, leeward.

Figure 14. Chordwise pressure distributions obtained in flight at specified angles of sideslip; Mach = 1.28,

Hp - 45,000 ft.

17
- 1.0

-- .6

-- .4

-- .2

Cp 0

-- .2

0

--.2

0

.2

- 1.0

-.6

.4

- .2

Cp 0

--.2

0

-- .2

0

.2

........ i...... i...... i-;,=o1.1

: : : -..-e-zfo'=0.48 |

......... -_......... 7......... _ -.- _': 0.86-1

0 .2 .4 .6 .8 1.0

x/c

960756

(b) 13- 1.93 °, leeward.

Cp

- 1.o i i i i ]

- .8 ......... -_......... " ......... !- --,-- Z/t)' : 0.16-

: : : .-=-- Z_'=0.48 |

_-".4......... _.......".........r-'-_'=°"-I

i :

-- .2 ............... +_........

0

.2

0

.2

0

•2 0 .2 .4 .6 .8 1.0

X/C

960757

(c) 13= -0.07 °.

/

........ ; ......... 4......... _-..-,- z_' = o.16 "'1

i i _ -.-_=o,m ]

........ _......... i......... :--.- ,_,'=o,, 1

........ i ......... i......... f......... 7........ i

.2 .4 .6 .8 1.0

_c

960758

(d) 13= -2.02 °, windward.

Figure 14. Concluded.

- 1.0

-- ,6

-- °6

-- .4

.2

Cp 0

_.2

0

- .2

0

.2

i i i ; I

........ "_......... ! ......... f-_ Z_' = 0.16 -t

......... _......... :......... :_.--.- z_' =o.48_j

I

0 .2 .4 .6 .8 1.0

_c

960759

(e) 13= -3.59 °, windward.

- 1.0

- .S ........... i.......... ]__ _ _ -' 8.04% leeward .__

, ' -4-- [_ = 3.95 °, leeward

.6

p

k ....... :.......... -:----#-- _ = o.oo....

_'l ! ! + 13= - 3.98 °, windward -4 ...._i .......... _--.-_=-..ooo. w,°°w.rO

Cp 0

f 'Nk' : l :

• .2

....

.4

,

.6

;i--"++.8 1.0

x/c

960760

(a) Mach = 0.52, Hp = 15,000 ft.

Figure 15. Chordwise pressure distributions obtained in flight at specified Mach numbers and angles of sideslip;

z/_' = 0.48.

18
Cp

- 1.0

-- .8

-- .6

-- .4

-- .2

0

0

0

0

0

.2

o

.__L.... L.... F.... _-_ _ = 4.01 , leeward -

kl 1 _,*, _13=-0.06 o

--,_ .... ,_.... _-._ _ = - 3.98 °, windwardi i t _ o

', _,, , , , --e-- I] = - 7.99 , windward

_----,,F----,,,'----'-- -" .... _-.... ,'-.... _-.... . .... . .... _-.... . .... , ,, ,, ,, ,. [ ', ,,

:-_-_--_ .... _.... ,_-_-

r;N. ; -', __;- .k . ± _r _

i "N_ i. k t t i i , A

i L i J J i

'. I ! '. '. i

, _=;= ,. -, :

i

i i * i ] i i i i

i * , = i i , ,

) L i * , p i

0 .2 .4 .6 .8

xlc

(b) Mach = 0.70, Hp - 45,000 ft.

1.0

960761

- 1.0

--,9

--,6

_,4

--.2

Cp 0

0

0

0

0

.2

i

t_;: ;I ---e-- [3 = 7.96 °, leeward

- _......... b- --m-- [3 = 3.99 °, leeward -

\, , --_ 13= - 0.03

......... :---.- [3--- w,ndwer 

(---\-k i .--e-- [3 = - 7.97 °, windward

..... : ........ _......... _......... _.........

0 .2 .4 .6 .8

x/c

(c) Mach = 0.89, Hp --- 45,000 ft.

Figure 15. Continued.

1.0

960762

19
Cp

-1°/ _ i--o--L_ = 3.57_ , leeward I

-.8 _-......... _......... _- --.m--_ = 1.93 , leeward -I

/ ! I

-.6 _-......... F......... _- --i-- p =-2.02 °, wlndward- i

| i i --e-- 13: - 3.59o,windward I

4 , ! , ,

u I i , i

0 .2 .4 .6 .8

x/c

(d) Mach = 1.28, Hp - 45,000 ft.

Figure 15. Concluded.

1.0

960763

offset --0.8, -0.6, -0.4, and --0.2 for 13of 3.6 °, 2.0 °, 0 °, and -2.0 °, respectively. Again, the most significant feature

of the chordwise pressure distributions as a function of Mach and j3 are the suction peak followed by recompression on the leeward side of the fixture. Another interesting feature is the change toward higher pressures near the aft

region of the F'IT-II at Mach 1.28 (fig. 15(d)).

Boundary-Layer Measurements

Because of the influence of the boundary-layer rake on the local static pressure at z/b' = 0.86 and x/c = 0.90,

velocity profiles (u/Ur) were calculated using a reference static pressure (Pw) orifice located at the same chord location and at z/b" = 0.48. Appendix D shows boundary-layer data obtained on the left side of the fixture tabulated for

the entire matrix of flight conditions. Local flow qualities such as the maximum local Mach number from the rake

(Mr), total airstream temperature (Tf), local wall temperature (Tw), and Pw have been included for steady-state

maneuvers to aid in planning experiments that may require further analysis of the boundary-layer. Reynolds numbers at the rake (based on a rake location ofx/c = 0.90) ranged from 8.02 × 106 to 36.9 x 106, depending on specific

flight conditions. For this report, boundary-layer thickness, 5, is defined as the distance from the wall where velocity

is 99 percent of the maximum local free-stream velocity. 12

The reference Mach number for the boundary-layer data is the M r obtained from the outer probes of the rake

that are outside the boundary layer. Figure 16 shows the relationship for the M r to the free-stream Mach number

obtained from the FTF-II noseboom, Moo. As figure 16 shows, little effect from altitude or Reynolds number exists

on the relationship of Mr/Moo. The M r shows a slight increase over the M._ at subsonic speeds. For M_ - 1.0 to 1.25,

the M r is slightly less than the Mo_ and then appears to increase to AM = 0.2 at greater than M., = 1.4.

Figures 17 and 18 show typical FTF-II boundary-layer profiles. The losses in the boundary layer at the respective altitudes are mostly consistent with the expected Reynolds number effect for the profiles shown (fig. 17). In general, large boundary-layer losses are observed at the low Reynolds numbers that occur at the high altitudes. Because

the boundary layer was not tripped at a fixed chord location, differences in the various profiles may have also resulted

from small differences in 13and transition location effects.

20
2.4

2.2

2.0

1.8

1.6

M r 1.4

1.2

1.0

.8

.6

.4

/

_ n _ n t t J t

o Hp = 15kft ..............

i i i i i i ___-- i °

[]Hp=3Okft • _ -, .... '-..... '- ..... ,..... -.... ._ .... ,_ , , ,

<>Hp ....... 45k ft _ /

.....

i t _ J i t n i

...... _...... i..... j ..... _ ..... L ..... r ............ A .....

i i t i i _ q i i

...... L_.... .J ..... .J ..... _ ..... L ............ -J..... _ .....

..... ,,...... . ..... _ ..... ; ..... ,,_..... ,,...... ,,..... q.....

.6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4

Moo

960788

Figure 16. In-flight relationship ofM r with Moo, outside the boundary layer, x/c = 0.90.

Y!

In,

2.5 ; ; ; ; ; ; '

I "-e-'HP = 15kft i i i '

/--- Hp ---30k ft ,i : , '

2 01--e- Hp -- 45k ft ..... _'..... "..... :..... _.....

1 5 i-.......... _................ ,-................

1.0

.5

0

.3

i i , l < I

i i i T

',1

.4 .5 .6 .7 .8 .9 1.0 1.1

u/U r

960764

Y!

in.

2.5 , i_

+ Hp = 30k ft

--e.-- Hp = 45k ft 2.0 -, ..... , ..... , .................

t

I

1.5 ..... ,_ ..... ,_..... T ..... +,..... ] ..... _ ..... _......

10.5 ...... ; ......................................

0

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

ulU r

960765

(a) Mach - 0.7. (b) Mach - 0.9.

Figure 17. Boundary-layer profiles obtained in flight at specified Mach numbers and altitude.

21
Yt

in.

2.5, . . ; . ; ; :

[ --e-- Hp - 30k ft : i : _ ,

I---_ Hp - 45k ft : : : :

L _ n i i i t i

i , i i i

20 ..... - ..... , ..... • ..... • ..... +..... " ..........

• l + i * q , i

, _ ) i i n

n I i i ) i

i i i i _ i

1 $ ..... , ..... "...... - .... " ..... " ..... " .... : .....

1.0 , , , _ , " .....

, : , , t

t i J ,

, j r i i

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

U/Ur 960766

(c) Mach -- 1.4.

Y+

in.

Figure 17. Concluded.

/ '='', ', ', ', ,t /

2.o_..... _..... i..... '_..... i..... i..... !..... 1-.....

i

1.5 ...... " ..... _..... - ..... - ........... ! ..... : .... ]

1°.5 iiiiiii iiiii t

0

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

u/U r

960767

(d) Mach - 1.5.

Yl

In.

2.5 ; ; ; ;

b

--"- Moo = 0.39 ,

Moo = 0.61

2.0 --- _ Moo:0.80

--*-- Moo = 0.98 ',

1.6........................ -,'..... _...........

1.0

.5

0

.3

: .... S

.4 .5 .6 .7 .8 .9 1.0 1.1

u/U r

960768

(a) Hp -- 15,000 ft.

2"6t i i i i i i "

----- Moo = 0.61 ', ', ; '

2.0 -- _ Moo=1.16 ---_ ..... ":..... _.....

Moo = 1.39 i i !

1.5 ---

Y,

in.

I

I

I......

° iiiiii .5 "

0

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

ulU r

9_0769

(b) Hp = 30,000 ft.

I --e-- Moo = 0.79

--e-- Moo = 1.16

2.0 _. _ Moo=1.60 ___, ..... , ..............

Moo = 1.80 _ "

Moo = 2.00 _ ,a

1.5 ....... - .... , ..... •..... • ..... • ..... _ V-- rd_......

in.

1.0 ............................. ; ..... _ ......

,5 ..... i .... _ ........,+

0

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

ulU r

960770

(c) Hp = 45,000 ft.

Figure 18. Boundary-layer profiles obtained in flight at specified altitude and Mach numbers.

22
Figure18showsboundary-layerprofilesat variousMachnumbersfor thethreeselectedaltitudes.At each

altitude,theprofilesfor Mach< 0.8 are generally similar with no significant features beyond mild inflections. The

profile at Mach 0.98 (fig. 18(a)) shows a more exaggerated inflection. The remaining profiles at supersonic conditions are quite different from those observed at the low Mach numbers. The presence of upstream shocks, located

uniquely for each Mach number, appear to impose unique upstream conditions on the boundary layer at the specific

Mach number. In addition, reinspection of surface pressure-distribution data (figs. 9 and 10) also shows the general

presence of an adverse pressure gradient in the rake location during supersonic conditions. Future boundary-layer

measurements forward ofx/c = 0.90 should be considered to identify and avoid any possible effects from an adverse

pressure gradient.

In general, boundary-layer thickness ranged from 1.25 in. to 1.45 in. with a slight increase at the supersonic

Mach numbers (figs. 18(a) and (b)). As an approximation, calculation of boundary-layer thickness for a flat plate

with a turbulent boundary layer 12using the origin of turbulent flow at x/c = 0.129 predicts a boundary-layer thickness

of 1.28 in. to 1.49 in. for the range of flight test Reynolds numbers. This prediction shows general agreement with

flight results. As a result, the following equation may be used to estimate boundary-layer thickness 12 for future

FTF-II experiments in a similar configuration with natural transition:

_/l = 0.37(Rel) -1/5

where l =

and x =

and Ax =

x-Ax, in.

distance from leading edge to test location, in.

12.84 in.

A more appropriate estimation for forced transition would be to set Ax as the location of the forced transition aft of

the FTF-II leading edge.

Figures 19 and 20 show boundary-layer profiles at various angles of attack and sideslip, respectively. Changes

in the boundary layer were observed as a function of changes in a (fig. 19). Large changes in the boundary layer, as

noted by the inflection in the boundary-layer profile, did not occur until tx = 4.05 °, which corresponds to t_A = 20 °.

2.5 , ; ; ; ;

! , • n a t

I I I "_ --e-- cx= - 1.85 ° ' ' '

o a i

--e-- a = 0.30 , , , _

2.0 ...... • ..... , ..... _.......... -rrl

a = 2.19 o

--*-- a = 4.05 o : ,

t _

J J J i _

p i i i i i

1.5 ........... , ..... • ..... • ..... , ..... , ........

Y, : : : : : :

t i i i t

In. _ , , , _ ,

i , , i t

1.0 ..... ,_..... , ..... ,*..... T..... _..... :-- ,......

i i a t I i t

i I u I q

I i n n I I

0 _ !

.3 .4 .5 .6 .7 .8 .9 1.0 1.1

u/Ur

960771

Figure 19. Boundary-layer profiles obtained in flight at specified FTF-II angles of attack; Mach = 0.54, 13= 0 °,

Hp = 30,000 ft.

23
Y|

In.

2.5 ; ; ; ; i ; _ :

--e-- 13= 8.04 o, leeward ' '

--.-- _ = 3.95 °, leeward

-- --_- I] = 0.00°

---4-- _ = - 3.98 °, windward

_ = - 8.00 o windward

r _- v ,r ............

i l

i i i

i i

i t

i i t

..... ; ....................... _.........

' '

i i i

t i n *

ol I t

.3 .4 .5

2.0

1.5

1.0

Y9

In.

1.0

2"51 ; • ; ; ; i ;_N[

I -'-- 13: 3.57o,I..werd : i _ _t

I --e-- _ = 1.93 °, INward : : _

2.0I- -.- _: - 0.070 -!..... _--\-_ ......

/ _ =-2-02°,windward : : "_1[

/ ---4--- 13=-3.59 °, windward i i

1.5 _..... " ..... :-..... ; ..... _..... 4..... " .... =,. ......

_ n t n

i _ i t t

i n i t

i n I n

i l n p n

i i i p t

! !

_ i I i i

• i i i i

' i : :

0

.6 .7 .8 .9 1.0 1.1 .3 .4 .5 .6 .7 .8 .9 1.0 1.1

u/U r u/U r

960772 960773

(a) Mach = 0.52, Hp = 15,000 ft. (b) Mach = 1.28, Hp = 45,000 ft.

Figure 20. Boundary-layer profiles obtained in flight at specified angles of sideslip.

This observation is a useful finding for experiments that plan to utilize a horizontal surface mounted to a fixture side

panel that may require slight changes in or.

Figure 20 shows boundary-layer profiles at sideslip conditions for subsonic and supersonic Mach numbers. At

Mach 0.52 (fig. 20(a)), the most distorted boundary-layer profiles are obtained for leeward flow at the sideslip angles

where 13> 3.95 °. At Mach 1.28 (fig. 20(b)), leeward flow also produces the most distorted profiles at 13= 3.57 °

Internal Environment

To characterize the internal environment within the FTF-II, bay temperatures were obtained on multiple flights

for a wide range of initial temperatures during preflight and for a variety of flight conditions to determine typical

and extreme temperatures. Acceleration data were also obtained to determine in-flight acceleration loads within

the FTF-II.

Bay Temperatures

For simplicity, the FTF-II has no active environmental controls to control internal temperature. As a result, bay

temperatures were monitored during all ground and flight testing of the FTF-II to document typical and worst-case

bay temperatures. Temperatures for bays A2, A3, T1, T2, T3, and T4 (fig. 4(a)) followed expected trends but were

dependent on ground and flight conditions. As expected, bay temperatures during pre- and post-flights were primarily affected by the ambient air temperature. In-flight bay temperatures were primarily affected by the ambient air

temperature at altitude and the duration of testing at altitude. Aerodynamic heating effects at the higher Mach numbers had little impact on temperatures within the fixture because of the limited duration of testing at these conditions.

Heating produced by the avionics or instrumentation within each bay affected individual bay temperatures, but the

trends in bay temperatures were consistent and easy to estimate for ground and flight operations.

Figure 21 shows bay temperatures for a typical flight profile with an ambient air temperature of 70 °F prior to

the flight. Bay temperatures ranged from 5 to 35 °F higher than the ambient air temperature, depending on the heating produced by the equipment in a specific bay. After aircraft takeoff, temperatures began to drop as expected. The

24
130

120

110

100

90

Bay

temperature, 80

OF

70

6O

5O

4O

B

/HHDMTake°ff =

= 4 = M 0.9,

- 4'5_'ft Hp - 45k ft /- Landing

'" -I M I ( D _ -=0.5, M -- 2.0, ,

_1 Hp = 15k ft I Hp = 45k ftl, _ Bay T1

I _---__ I t J i

I _kl _ I r Bay A3 I I

-I i t f ./i

_ , ,_./ ,

- i l

_ ! J

I I I

I i J I l II i II I

3O

0 15 30 45 60 75 90 105

Time after takeoff, min

960774

Figure 21. FFF-II bay temperatures obtained during a typical flight profile.

lowest observed bay temperature of approximately 30 °F occurred in bay T4 after approximately 70 min of flight

testing at altitudes between 15,000 ft and 45,000 ft. Temperatures began to increase with a return to low altitudes

prior to landing at the completion of the flight. From these data, the maximum expected range of internal temperatures within the FTF-II is 30--105 °E Because ambient air temperatures at NASA Dryden can vary +40 °F, projected

worst-case internal temperatures within the FTF-II would be -10 °F in-flight to 145 °F during ground operations.

More extreme temperatures could exist, depending on a wide variety of experiment variables. Experiments with

temperature sensitive equipment are advised to use thermal control systems or monitor bay temperatures near the

location of the flight experiment.

Vibration Data

High-frequency data from piezoelectric accelerometers were obtained to document local three-axis accelerations inside the FYF-II. Accelerometer data were recorded on tape at various flight conditions, low-pass filtered at

2000 Hz, and digitized at 8192 samples/sec for further analysis. Short time segments (1 sec) of relatively constant

flight conditions were used to calculate root mean square (rms) magnitudes and power spectral density (PSD)

functions. Each PSD is an average of 30 overlapped records spanning a given 1-sec interval.

During the analysis process, y-axis data were determined to be unreliable. As a result, only x- and z-axis data

have been summarized in this report. Figure 22 shows acceleration and PSD results from typical flight maneuvers

for the x and z axis. At Mach 0.70 (fig. 22(a)), g levels tend to increase with [3.At [3-- -5 °, maximum grms is approximately 6.0 and 3.5 for the x and z axes, respectively. The corresponding PSD plot indicates that most the accelerations occur as random input at greater than 200 Hz. For clarity, only the maximum and minimum values for a given

PSD set are plotted. Data from an acceleration/deceleration between Mach 1.0 and 1.5 and from Mach 1.9 to 2.0

show similar grms levels and spectral content (fig. 22(b)). Figure 22(b) also shows that g levels tend to increase with

Mach number. At Mach 2.0, maximum grms is approximately 6.0 and 3.0 for the x axis and z axis, respectively.

25
gtme 4

g2/Hz

2

10-2

10-4

10-6

0

-6

100 --

/-- x axis

/ x

x

x

x

X ×

I I I I I

-4 -2 0 2 4

_,deg

Curve B -_

- A

\ , _

_" "-z'axJs, me .....__" _

z axis, minimum j

I

101

I I I I I I III I I I 1 I I III

lo2 lOa

Frequency, Hz

(a) Sideslip, Mach = 0.70, Hp = 15,000 ft.

Figure 22. FTF-II accelerometer data obtained in flight.

960775

26
x axis

grms 4

Z aXle

g2/llz

C

1.0

t I

1.2 1.4

Mach

I

1.6

I

1.8

I

2.0

100 B

10-2

10-4

10-6

101

C U rveL,.'-_,_.._

..... __-e_"; _''" "' , _ L_i

z°xJ.,m.imum-/ /_ _t_

. x axis' minlmum ,_'_ _ ili'_;ill _/__f'_'_'_"_ _I

10 2 103

Frequency, Hz

960776

(b) Acceleration/deceleration, Hp = 45,000 ft, _ = 0 °.

Figure 22. Concluded.

27
Thesignificantresultfromanalysisof theaccelerometer data(x andz axis) is that grms < 6.0 for most flight

conditions within the operational envelope of the F- 15B/FTF-II. Further details on how the flight data can be related

to FTF-II vibration test requirements, such as Curve B (fig. 22), are included in the next section of this paper. A

more comprehensive study, including y-axis data, of FTF-II local accelerations must be completed before a

reduction in the FTF-II vibration test requirements can be considered for all experiment-related hardware.

Designing Fixture Experiments

Understanding the aerodynamics and internal environment of the FTF-II is important but is only a contributing

factor in the development of a flight experiment. Successful design of an FTF-II experiment will require a basic understanding of established NASA Dryden policies and guidelines as well as alternate FTF-II configurations that can

be achieved to meet specific experiment requirements. This section serves as a guide to assist in the development of

future experiments on the FFF-II. Additional information enclosed in the appendices has been included to assist

with more detailed analyses by project personnel, if required.

Established Policies and Guidelines

NASA Dryden requires flight projects such as the F-15B/FTF-II to follow established policies and guidelines

to help achieve efficient flight operations without compromising safety. The most relevant NASA Dryden documents

for developing flight experiments on the FTF-II are the Basic Operations Manual, Procedures for Objectives &

Research Requirements Documents, t and Process Specification Manual. _"All of these documents are available to

NASA Dryden personnel, and specific requests can be accommodated. The Basic Operations Manual outlines the

basic established policies for NASA Dryden flight projects and provides some guidelines to assist with implementing these policies. The Basic Operations Manual reviews policies on issues such as risk management, system safety,

quality assurance, project plans, configuration management, flight preparation, and flight readiness and review

processes at NASA Dryden.

Much of the input into project plans is derived from information that is submitted in the form of an Objectives

and Research Document (ORD). An ORD is the link between the technical objective and the proposed experiment

requirements. All F-15B/FTF-II experiments require that an ORD be submitted to the project for evaluation and approval. As a minimum, this ORD should include 16 an experiment title, an introduction, an objective, the test scope

and approach, specific requirements, and the appropriate approvals. Early development and approval of an ORD ensures that specific experiment requirements are defined and reasonable flight objectives can be achieved in a safe

and timely manner by the project.

Another key document for compliance in the development of a flight experiment is the NASA Dryden Process

Specification Manual. The purpose of this document is to provide specification for the acceptance of avionics, instrumentation, or experimental hardware mounted internally or externally on the F'rF-II. The primary specification

sections of interest for potential flight experiments are "Section 00 - Operations" and "Section 21 - Testing." For

testing, because the F-15B airplane is a jet aircraft and the FTF-II is installed at the centerline pylon location, all

FTF-II hardware must meet the vibration test requirements for Category II aircraft Curve B. Because the FTF-II is

not a temperature-controlled environment, all electronics should be tested for operation from -65 to 160 °F and to

*National Aeronautics and Space Adrmnistration Dryden Flight Research Center Basic Operations Manual, X- 1.000, Rev. 5, Feb. 1995.

?Bauer, Carol A., Procedures for Objectives & Research Requirements Documents (ORDS), Dryden Flight Research Center, Rev.

12/87, Dec. 1987.

:_NASA Dryden Flight Research Center, Process Specification Manual, Rev. 5/95, May 1995.

28
anabsolutepressureequivalentof 60,000ft, whichisthealtitudelimit of theF-15Baircraft.Vibrationtestsonall

hardwareshouldbeperformedusingarandomvibrationtestcurveequivalentto the12.2g2/Hz, to a maximum

2,000 Hz (fig. 22). These test specifications are the result of many years of flight test experience and should be

followed whenever possible to ensure flight safety, avoid aircraft operating restrictions, and ensure a successful

flight experiment.

Alternate Fixture Configurations

Past experience with theF-104 FTF has shown the value of a versatile fixture where configuration changes can

be accomplished without major structural changes to the fixture. A major design goal during the development of the

FTF-II was to maintain the maximum practical configuration versatility. As discussed previously, the primary

FTF-II shape (fig. 4) can be modified to satisfy the requirements for a variety of flight experiments. The lower 13 in.

of the FTF-II, the vertical test article, can be replaced by aerodynamic shapes that may provide a better match to

specific experiment requirements. Other configuration changes (fig. 4) can range from alternate side panels to an

alternate nose section for the avionics pylon and test article.

Several "standard" alternate configuration changes have already been developed for the FFF-II. Spare nose

section test panels and test article aft left side panels have been fabricated for future experiments. An aft fairing that

closely matches the nose section contour (fig. 23) was installed after obtaining the pressure data presented in this

paper in order to improve the aft region flow quality, reduce the base drag of the FTF-II, and reduce the resulting

base area effects that cause the pressure gradients observed in the pressure data at x/c > 0.90. An example of a significant configuration change to the FTF-II was for the thermal protection systems durability program (fig. 24). For

this program, an alternate nose section was installed to contain tile specimens, and the vertical test article was

replaced by a particle measurement probe.

Figure 23. FTF-II with aft fairing and flow visualization results obtained in flight, right side with tufts; Mach = 0.70,

Hp = 45,000 ft.

29
Figure24.FTF-IIconfigurationforthethermalprotectionsystemsdurabilityprogram.

AnimportantupgradetotheF-15Bairplanewastheinstallationof on-boardvideocapabilitytodocumentflight

experiments. Thecurrentvideocoverage(aftleft andfrontright videocameras)(fig.25) shouldbesufficientfor

mostflightexperiments. Althoughthecurrentvideocoverageiscomprehensive, videocameralocationscanbemodifiedtoaccommodatespecificexperimentrequirementslikethefrontleftvideocamera(fig.25).

Eachconfigurationchangewill requireanevaluationwithrespecttothefixturestructureandexperimental hardware.Typically,thebestapproachisto "over-design"experimental hardwaresothatflightsafetyissuescanberesolvedusingsimpleandconservativeengineeringanalyses. Configurationchangessignificantlydifferentfromthe

_ View from eft left _" _ _- --... _._

_ video camera "_-._.._ ................

_ "" __.=_..._......... View from front /

.,4 ..... "_ ___ left video camera /

, ¢

""- - "- - View from front

"'---'---. right video camera _

960780

Figure 25. F-15B on-board video camera locations, top view.

30
currentfin-likeshapeandareamayrequireadditionalstructuralloadsoraircraftstabilityanalysesasappropriate.

AppendicesA andBcontainfurtherdiscussionof structuralloadsandaircraftstabilityconsiderations, respectively.

SUMMARY OF RESULTS

NASA Dryden Flight Research Center has developed the Flight Test Fixture II (FTF-II) for use as a generic test

bed for aerodynamic and fluid mechanics research. The FFF-II is a low-aspect-ratio vertical fin-like shape that is

mounted on the centerline of the F- 15B lower fuselage and has flown at Mach numbers to a maximum of 2.0. The

FI'F-II is a composite structure with a modular configuration and removable components for functional flexibility.

Flow-quality data were primarily taken at 1-g flight conditions and a matrix of nominal altitudes of 15,000 ft,

30,000 ft, and 45,000 ft throughout the airspeed envelope of the aircraft. The flow environment was characterized

using FTF-II noseboom airdata combined with surface pressure and boundary-layer measurements to survey aerodynamic flow quality. The FTF-II angle of attack at the trim aircraft angle of attack ranged from 1o at subsonic

speeds to -9 ° at Mach 2.0. In general, chordwise pressure gradients were small aft of the nose section suction peak

and recompression, except for the presence of a weak shock in the forward region of the FTF-II and base area effects.

Boundary-layer thickness at the chordwise location x/c = 0.90 ranged from 1.25 in. to 1.45 in. with a slight increase

in boundary-layer thickness at the supersonic Mach numbers.

Environmental conditions within the FTF-II were also documented to determine typical and extreme testing

conditions. Projected worst-case internal temperatures within the FTF-II are -10 °F in flight to 145 °F during ground

operations. The analysis of accelerometer data (x and z axis) indicates that grms < 6.0 for most flight conditions

within the operational envelope of the F-15B/FTF-II.

Successful design of an FI'F-II experiment by future investigators will require a basic understanding of established NASA Dryden policies and guidelines and alternate FrF-II configurations. The most relevant NASA Dryden

documents for developing flight experiments on the FTF-II are the Basic Operations Manual, Procedures for

Objectives & Research Requirements Documents, and Process Specification Manual. Alternate FTF-II configurations can be developed to satisfy a variety of flight test requirements, but these configurations must consider related

operational and flight safety issues.

The FTF-II is a proven and documented flight test facility for aerodynamic and fluid mechanics research at

speeds to a maximum Mach 2.0. Improvements to the F-15B performance, combined with a new FTF-II nose section

and some additional analysis and verification of FrF-II loads and aircraft stability, could extend this limit as high

as Mach 2.3.

ACKNOWLEDGMENTS

The author would like the thank the following contributors who provided valuable expertise and guidance toward

this effort: Edwin J. Saltzman in the discipline of aerodynamics, Robert W. Kempei for issues related to aircraft

stability, and Lawrence C. Freudinger in the reduction and analysis of acceleration data.

31
REFERENCES

1 Meyer, Robert R., Jr., A Unique Flight Test Facility: Description and Results, NASA TM-84900, Nov. 1982.

2 Montoya, Lawrence C., Brauns, David A., and Cissell, Ralph E., Flight Experience With a Pivoting Traversing Boundary-Layer Probe, NASA TM-X-56022, Jan. 1974.

3 Chiles, Harry R. and Johnson, J. Blair, Development of a Temperature-Compensated Hot-Film Anemometer

System for Boundary-Layer Transition Detection on High-Performance Aircraft, NASA TM-86732, Aug. 1985.

4 Moes, Timothy R. and Meyer, Robert R., Jr., In-Flight Investigation of Shuttle Tile Pressure Orifice Installations, NASA TM-4219, Sept. 1990.

5 Saltzman, Edwin J. and Hintz, John, Flight Evaluation of Splitter-Plate Effectiveness in Reducing Base Drag

at Mach Numbers From 0.65 to 0.90, NASA TM-X-1376, May 1967.

6 Trujillo, Bianca M. and Meyer, Robert R., Jr., In-Flight Load Testing of Advanced Shuttle Thermal Protection

Systems, NASA TM-86024, Dec. 1983.

7 Meyer, Robert. R., Jr. and Barneburg, Jack, In-Flight Rain Damage Tests of the Shuttle Thermal Protection

System, NASA TM-100438, May 1988.

8 Zuniga, Fanny A., Anderson, Bianca T., and Bertelrud, Arild, Flight Test Results of Riblets at Supersonic

Speeds, NASA TM-4387, June 1992.

9 Richwine, David M. and Del Frate, John H., Development of a Low-Aspect Ratio Fin For Flight Research

Experiments, NASA TM-4596, Aug. 1994.

10Tanaka, Arthur Y. and Huete, Rodrigo J., "F/TF-15A Flying Qualities Air Force Development Test and Evaluation," AFFTC-TR-76-48, July 1977.

11 United States Air Force, "Flight Manual: USAF Series F-15A/B/C/D Aircraft, Block 7 and Up," TO 1F15A- 1, July 1989.

12 Schlichting, Hermann, Boundary-Layer Theory, McGraw-Hill, Inc. New York, 1979.

13 Shapiro, Ascher H., The Dynamics and Thermodynamics of Compressible Fluid Flow, The Ronald Press

Company, New York, 1953.

14 Anderson, John D., Jr., Hypersonic and High Temperature Gas Dynamics, McGraw-Hill Book Company,

New York, 1989.

32
APPENDIX A

STRUCTURAL LOAD CONSIDERATIONS

FOR ALTERNATE CONFIGURATIONS

Major configuration changes to the Flight Test Fixture II (FTF-II) could have an impact on FTF-II loads and

the F-15B airplane, depending on a variety of factors. The four load cells located at the sway braces (fig. 4(b)) can

be used to validate side loads for rapid envelope expansion of major configuration changes. When needed, in-flight

loads can be monitored and compared to prior loads documented in this report for the baseline fin-like configuration.

Flat-plate approximations were used to estimate FTF-II side loads in terms of aircraft limits, where side force,

Fy, is approximated by (sin _)*qo_*As.9 Ground tests of the FTF-II mounted to the F-15B pylon (fig. A-I) were then

conducted to correlate known applied side loads on the fixture to the measured resultant loads at the sway braces.

By subtracting initial preloads (500-1000 lbf) at each sway pad and combining sway pad loads for each side, sway

pad reaction loads, Fsw, were related directly to the applied moment, Mx, on the fixture. Figure A-2 shows this

relationship was quite predictable and repeatable over several load cycles, where Fsw = 39.4 + 0.385"M x.

Initial load assumptions, estimates, and factors of safety for the FTF-II design were quite conservative. 9 As a

result, a design limit equivalent to M x = 48,368 in-lbf (combined inertial and aerodynamic side loads 9) was used for

analysis of the fixture. Initial estimates were that an M x equal to 48,368 in-lbf would result in an Fsw equal to

23,884 lbf. Determining actual loads using a linear extrapolation of ground test data indicates that the design limit

Mx = 48,368 in-lbf would generate Fsw equal to 18,661 lbf. The difference between estimated and actual reaction

loads is attributed to the indeterminate nature of the suspension lug/sway brace mounting structure. This result was

also observed in the initial loads analysis and ground testing of a fixture test article. 9 The relevant aspect of this evaluation with respect to a flight experiment is that in-flight side loads that exceed the design limit of the FTF-II will

indicate Fsw > 18,661 lbf.

Figure A-1. Load testing of the FrF-II mounted to the F-15B pylon.

960782

33
M x ,

in-lbf

15,000

10,000

5,000

---4-- Baseline from ground calibration

o Flat-plate approximation, flight o

o Surface pressures, flight o

0

o 0 •

D

O0

0

0

0 aid 0 FSW = " " *

t

I I 1 I I

0 1000 2000 3000 4000 5000

Fsw,lbf

960783

Figure A-2. The FTF-II reaction loads at the sway braces, flight- and ground-test data.

Further in-flight load evaluations were performed to gain additional confidence in estimated and actual loads.

Integrated surface pressures and the flat-plate approximation using z/b' = 0.50 were calculated for the matrix of flight

conditions. These techniques (fig. A-2) show reasonable agreement with each other but tend to overestimate flight

loads by more than 50 percent for a given reaction load at the load cells. At the same time, flight data (fig. A-2) also

reveal that existing F-15B control system limits, as observed from flight data, limit aircraft aerodynamic loads to

less than 30 percent of FTF-II design limit side loads. These data imply that aerodynamic loads for any reasonable

FTF-II configuration are fairly insignificant with respect to exceeding the design load. The primary area of future

concern would be to avoid very dynamic maneuvers where high inertial loads in the lateral axis can be coupled with

aerodynamic loads. As a simple and conservative rule, inspection of the FTF-II should be conducted if the total

Fsw > 10,000 lbf (greater than 53.6 percent of FrF-II design limit side load) during flight test.

Major configuration changes to the FTF-II could also have an impact on the FIT-II center of pressure and the

resulting distribution of loads on the fixture. Initial side force estimations for the baseline fin-like configuration 9

used a center of pressure of x/c = 0.15 to x/c = 0.66 for the estimation of aerodynamic loading. Further in-flight load

evaluations at the matrix of flight conditions were performed using integrated surface pressures to document the

actual FFF-II center of pressure.

The flight data (fig. A-3) indicate that the center of pressure estimations were valid, over a range in Mach

numbers of 0.50 to 1.28, for an expected range of x/c = 0.15 to 0.66. 9 At subsonic speeds, the center of pressure

ranges between 0.15 < x/c < 0.40 with the farther aft locations at sideslip conditions where [3 > 6 °. Limited sideslip

data at Mach 1.28 show the general trend for the center of pressure to move slightly aft at supersonic speeds as well.

Extrapolation of the flight results for the baseline fin-like configuration shows that development of a flight experiment requires only minimal consideration for the center of pressure location. In general, only major configuration

changes or extensions of the nose section that would move the overall FTF-II center of pressure forward of the

current x/c = 0.15 location would require further analysis.

34
1.4

1.3

1.2

1.1

1.0

Mach .9 number

.8

m o z/b' = 0.16

[] z/b' = 0.48

-- 0 z/b' = 0.86

• 7 B

,6

.5

.4 I

0 .1

0

Expected range for

]dc (subsonic with 13= 0°)

(1)

_0 o

I I

.2 .3

xlc

x/c increases as:

(1) Supersonlc

(2) 13increases

di>'_P (2) _4°

I I

.4 .5

960784

Figure A-3. Center of pressure locations determined from flight data as a function Mach number.

35
APPENDIX B

AIRCRAFT STABILITY CONSIDERATIONS

FOR ALTERNATE CONFIGURATIONS

Major configuration changes to the FTF-II could have an impact on aircraft stability, depending on the size and

location of the configuration change. Initial envelope expansion of the F-15B/FTF-II configuration (vertical fin)

consisted of flight test of the FTF-II uninstrumented on the F-15B aircraft to perform envelope expansion 9 to a maximum of Mach 1.3 in the vertical fin configuration. Directional stability during phase one flight tests was evaluated

by performing rudder doublets and making qualitative comparisons to the F-15B aircraft with and without a centerline tank. Comparisons at subsonic speeds indicated that the subsonic static directional stability, Cnf_, derivative levels for the F-15B/P'TF-II configuration were greater than 0.002/deg and had no adverse effects on aircraft

stability. 9,10 Predictions also indicated that the Cnf_ would decrease with increasing supersonic Mach numbers and

approach levels as low as approximately 0.0005/deg at Mach > 1.5. During phase two envelope expansion of the

F-15B/FTF-U to Mach 1.8, several actions were taken by the project to understand and minimize the potential risk

of reduced Cn_ caused by positioning the FIT-II at the centerline of the F-15B aircraft.

A study was performed to assess the risk of F-15 control augmentation system (CAS) disengagements and their

impact on flight safety at reduced Cnfy The CAS disengagements were found to occur approximately once in every

200 flight hours and mostly during dynamic maneuvering. The CAS disengagements produced transients that were

easily tolerated by the pilot and aircraft. In the case of a CAS failure, a 600-KCAS airspeed limit is imposed because

of a mildly divergent longitudinal stability. In addition, a Mach 2.0 limit exists for the two-seat F- 15 aircraft because

pilots comments were that uncomfortable amounts of sideslip are produced. Overall, CAS disengagements at high

speeds were found to result in handling qualities undesirable for sustained operations but quite acceptable for safe

aircraft maneuvering.

An analysis of F-15 lateral--directional stability without the CAS, using a NASA Dryden batch simulation, was

also performed for airspeeds to a maximum of Mach 2.2 at an altitude of 45,000 ft with particular attention to the

dutch roll mode. Three linear models of an F-15 aircraft at a gross weight of 36,200 Ibm were analyzed: one with

predicted lateral-directional aerodynamics, one with aerodynamic increments caused by the lZTF-II added, and one

with reduced increments to match Air Force flight test data. 10 From these models, the damping ratio, 4, and undamped natural frequency, mn, were calculated for the dutch roll mode and compared to MIL-F-8785C for Class IV

aircraft (fighter) category B (cruise) handling qualities requirements. As figure B-1 shows, the addition of the FI'F-II

to the F-15B aircraft improved dutch roll damping at an altitude of 30,000 ft with sufficient stability at all flight

conditions as predicted. Data at higher altitudes were also calculated and included with comparisons to flight

test data.

F-15B/FTF-II envelope expansion was performed with rudder doublets to a maximum of Mach 1.8 with the

CAS on and off for pilot evaluations and comparison to model predictions. Satisfactory-to-good damping was reported by the pilots with the CAS on or off at airspeeds to a maximum of Mach 1.8. Acceptable, but low-tomoderate, damping was noted at Mach> 1.5. With the CAS off at Mach > 1.5, pilot comments concurred with predictions of handling qualities undesirable for sustained operations but quite acceptable for safe aircraft recovery.

Flight data from the rudder doublets were also analyzed to calculate _*c0n for the dutch roll mode for comparison

to predicted values (without the FI'F-II, CAS off) and MIL-F-8785C (fig. B-2). Figure B-2 shows the flight data

with the FTF-II installed and the CAS off. The F-15 airplane with the FTF-II installed appears to be more stable

than predictions of an F-15 aircraft without P-TF-II. As expected, _*e0n decreases with increasing altitude. At

Mach 1.8 and an altitude of 45,000 ft, _*mn is approximately 0.25 rad/sec, which is in the lower region of Level 3

handling qualities.

36
Thefinal phaseof F-15B/FTF-IIenvelopexpansionwasperformedwithrudderdoubletsto a maximumof

Mach2.0withtheCASon.ManeuverswiththeCASoffwerenotattemptedat speedsgreaterthanMach1.8.No

largereductionsin stabilitywereobserved,andpilotcommentsindicatedthathandlingqualitieswereacceptablefor

safeaircraftmaneuvering. PredictedtrendsindicatethattheF-15/F'IT-IIdutchroll modewiththeCASonoroff is

stableatMach_<2.2.Additionalevaluationof aircraftstabilityby performingrudderdoubletswiththeCASonis

recommendedfor significantchangesto theFTF-II or if envelopeexpansionof thecurrentFTF-II configuration

beyondMach2.0atanaltitudeof 45,000ft isdesired.

/

' ; : : : : : ' ' :/ ; :

: , , __ , I, / ,FTF-Ilon--_ , ,

--H .... :.... _- -,--,- - -"-T'-_-_-- - :-- - - : .... ;--A_ .... :---- -'-.... ,....

i ! /"_/MI'L-F-8785C i ! i//_ !

...: .... __..:. Level2 --;---:---_'--,;,,_-TL---+ .... "---

.5 : : : handling : i /.,,_.' .,,"Z.' i ! :

: : : qualities : _...j1- ', : : :

"i ....::.... ....i....

•4 "'-i .... _iTL ! --I_;e; 3 i .... i---i .... i

_* tOn .... i ---L- i i : /, handling : : : _.

" , --, .... , -"" T .... '- -1- '_- - qualities ..... '_.... ?-; rad/iec ....... ..

' ' .- __ ' .)/ : (Category B) :. -" :

.3 ........ ,--.,"f.... ,......... ,.... r---7.... ,----i .... ,----'T.... ,....

........ "_.." .... ;.......... :............. :.._.,;K ...............

, , ', ; ,_,,: : ,,:1 , ', ,, :

.... :.... _.... :......... :.... __.._,._,,_____._ .... L............. .2 .... i i i i i..": i i ! :

---" .... ;-- -- ;-" -" _ :- :, -- --',_ - - ,_- - - _,.... ,_--I .... _--; .... _---

1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4

Mach number

960?85

Figure B-1. The F-15B dutch roll mode damping obtained from NASA Dryden batch simulation (GAS off) with and

without FTF-II; Hp = 30,000 feet, aircraft weight of 36,200 Ibm.

i i ! , i/i i i/ i i i

---i........ i:,-,.-.'------_--'-¥---_---_ .... i--/ .... "---i .... i....

' ' ' ', MIL-F-8785C ', / : ' ' '

' /'_ ' . ' 30kft

: ,/,,_',.,. ;._ :Level2handhng :/ : !j,_, ,

.5 ---_ .... ;--,_v.,_,[-./_---T-- qualities --J --- _. --- _ .... Y---

i i i / i i (Category B)/'.j_ Level 3 i

---! .... !-'-]_---! .... t .... !.... !-7_'-- !" handling "-!....

', ', /_ !/._ i : qualities ',

..... _' ..... 4

t

tOn

.......... L__.: .... ;.._., .... L._[k _ • .... L---*' .... '_- ---_- L0-k -/1--,,

rad/sec ...... .' ...... -

_ "f__. _:....

: :"__--o_i ;/i :

, _ .... _..__.........,____........ :.-___,__..: ,

---_ ...... • .... _- --_ .... ,.... %;,.;.;.1, .... _ ---p .... _-- • ---_---

, , 4 ',J

, , , _ q , Jq

.2 .... _.... L.__,.... ,.... _.... .... :___Jt :

; : : : .. _" - - : -.-0-- Flight, 30k/1

.... i .... ;---; .... _ - - - ; - - - -',_-_'- _-- - -_.... _. '--I_--" Flight, 45k ft

..... Dryden batch

" : " : : _ ' , ', limulaUon

1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4

Mach number

960786

Figure B-2. Comparison of predicted and flight test values for F-15B dutch roll mode damping, CAS off, FI'F-II

on for flight, FTF-II on for simulation.

37
APPENDIX C

CHORDWISE PRESSURE DISTRIBUTION DATA

Table C-1. Steady-state chotdwise pressure distribution data obtained in flight, Hp - 15,000 ft.

F.r_-$tream Mach number

z/b" x/c 0.39 0.52 0.61 0.70 0.80 0.90 0.98

0.16 0.000 1.010 0.990 1.040 1.095 IAI4 1.204 0.980

0.025 -0.102 -0.112 -0.193 -0.248 -0.222 -0.108 0.000

0.050 -0183 -0.195 --0,285 --0.361 -0.370 -0.298 -0.169

0.075 -0,197 -0.210 -0.305 -0,388 -0,413 -0,365 -0,241

0.I00 -0.216 -0.227 -0.320 -0.404 -0.437 -0.416 -0.290

0.125 -0.202 --0.206 -0.299 -0.383 -0.413 -0,415 -0.297

0.150 -0.154 -0.165 -0.253 -0.333 -0.360 -0.483 --0.354

0.210 -0.086 --0.102 -0.184 -0.255 -0.274 -0,290 -0.203

0.260 -0.113 -0.131 -0,217 -0.302 -0.353 -0.303 -0.335

0.320 -0.088 --0.I01 -0.179 -0.257 -0.306 -0.389 -0.357

0.370 -0.072 -0.081 -0.155 -0.218 -0.24.4 -0.481 -0.410

0.420 -0.055 -0.069 -0.137 -0.197 -0,214 -0.410 -0,381

0.470 -0.065 -0.070 -0.139 -0.195 -0.212 -0.240 -0.422

0.540 -0.053 -0.061 -0.126 -0.181 -0.195 -0.182 -0.346

0.600 -0.036 -0.037 -0.102 -0.149 -0.151 -0.132 -0.220

0.700 -0.019 -0,018 -0.075 -0.116 -0.117 -0.093 -0.115

0.800 -0035 -0.029 -0.084 -0.125 -0.126 -0.100 -0.111

0.900 -0048 -0.040 -0.091 -0.131 -0.127 -0.091 -0.069

0960 -0.180 -0.178 -0.228 -0,318 -0.377 -0.282 -0.212

0.48 0.000 10l 9 1.007 1,040 1.096 I. 111 1.203 0.930

0.025 -0.119 -0,126 -0.209 -0.261 -0.225 -0.089 0.010

0.050 -0.225 -0.234 -0.330 -0,417 -0.427 -0.327 -0.205

0.075 -0242 -0.259 -0.366 -0.471 -0,524 -0.460 -0.329

0.I00 -0.248 -0.26,1 -0371 -0.477 -0.544 -0.576 -0.438

O. 125 -0.257 -0.265 -0.367 -0.464 -0.514 -0.656 -0.516

0.150 -0.196 -0.205 -0.293 --0.377 -0399 -0.654 -0.535

0.210 -0.088 -0.096 -0.168 --0.228 -0.243 -0.123 -0.341

0.260 -0.061 -0.069 -0.139 -0.202 -0230 -0.173 --0.173

0.320 -0.084 -0.079 -0.139 -0.198 -0.222 -0.277 -0.269

0.370 -0.043 -0.047 -0.114 -0.173 -0.192 -0.352 -0.270

0.420 -0.037 -0.041 -0.104 -0.156 -0.173 -0.361 -0.338

0,470 -0036 -0,037 -0.096 -0.150 -0.162 -0.144 -0.329

0540 -0,031 -0.033 -0.094 -0.141 -0.154 --0.143 --0.381

0.600 -0.030 -0.026 -0.081 -0.127 -0.131 -.0.115 -0.221

0.700 -0.030 -0.022 -0.077 -0.117 -0.117 -0.099 -0.133

0800 -0.032 -0.027 -0.077 -0.117 -0. t17 -0.094 -0.120

0.900 -0.063 -0.058 -0.112 -0.152 -0.149 -0.106 -0.079

0.960 -0106 -0.101 -0.167 -0.223 -0.239 -0.189 -0.185

0.86 0.000 0.980 0.970 1.010 1,062 1.074 1.157 0910

0,025 -0109 -0.131 -0.224 -0.300 -0.296 -0.182 -0.057

0.050 -0.195 -0.213 -0.312 -0.409 --0452 -0.425 -0.287

0.075 -0.212 -0.227 -0.327 -0.426 -0.485 -0.482 -0356

0.100 -0.210 -0.229 -0,325 -0.423 -0487 -0.560 -0433

0.125 -0.195 -0.197 -0.286 -0.375 -0.407 -0.611 -0.481

0.150 -0 153 -0.169 -0.254 -0.328 -0.351 -0.647 -0.523

0.210 4) 056 -0.065 -0.135 -0.190 -0.203 -0,084 -0.323

0.260 -0040 -0.043 -0.108 -0.160 -0.177 -0.143 -0117

0,320 -0.018 -0.028 -0.094 -0.147 -0.164 -0.206 -0.151

0.370 -0037 -0.037 -0.093 -0,145 -0.159 -0.242 -0203

0420 -0027 -.0.023 -0.084 -0.132 -0.144 -0.271 -0.208

0470 -0.023 -0.019 -0.076 -0.124 --0137 -0.146 -0.268

0.540 -0.020 -0.020 -0,078 -0.126 -0.133 -0.128 -0.276

0.600 -0.021 -0.016 -0.070 -0.113 -0120 -0.111 -0.306

0700 -0.019 -0.016 -0.075 -0117 -0.123 -0.110 -0.144

0.800 -0.023 -0.016 -0.073 -0.113 -0.114 -0.096 -0.108

0.900 -0.027 -0,027 -0.077 -0.107 -0,089 -0.010 0.020

0960 4).148 -0.125 -0.185 -0.231 -0.228 -0.177 -0.151

38
Table C-2. Steady-state chordwise pressure distribution data obtained ia flight, Hp _ 30,000 ft.

Free-stream Mach number

z/b" x/c 0.51 0.62 0.70 0.79 0.89 1.00 1.05 1.15 1.28 1.39 1.48

0.16 0.000 1.005 1.045 1.092 1.132 1.207 1.202 1.246 1.330 1.307 1.149 !.045

0.025 --0.075 --0.162 -0.190 --0.186 -0.091 0.020 0.090 0.200 0.280 0.250 0.280

0.050 -0.172 -0.265 -0.315 -0.339 -0.277 -0.153 --0.098 -0.002 0.090 0.030 0.070

0.075 -0.183 -0.282 -0.342 -0.378 -0.335 -0.211 -0.156 -0.090 0.000 -0.072 -0.040

0.100 -0.196 -0.301 --0.357 -0.398 -0.394 -0.263 -0.206 -0.158 -0.071 -0.167 -0.123

0.125 -0.181 -0.279 -0.343 -0.379 -0.391 -0.269 -0.204 -0.147 -0.061 -0.218 -0.187

0.150 -0.136 -0.232 -0.289 -0.323 -0.448 -0.316 -0.243 -0.163 -0.065 -0.138 -0.229

0.210 -0.070 -0.161 -0.215 -0.243 -0.172 -0.168 -0.097 -0.010 0.090 0.090 0.050

0.260 -0. i 05 -0.200 -0.264 -0.313 -0.280 -0.311 -0.258 -0.204 -0. ! 00 -0.051 -0.042

0.320 -0.075 -0.166 -0.225 -0.262 -0.369 -0.360 -0.315 -0.309 -0.213 -0.153 -0.113

0.370 -0.057 -0.133 -0.183 -0.211 -0.289 -0.387 -0.312 -0.291 -0.199 -0.167 --0.127

0.420 -0.047 -0.123 -0.169 -0.186 -0.154 -0.350 -0.308 -0.265 -0.234 -0.227 -0.182

0.470 -0.050 -0.124 -0.169 -0.184 -0.156 --0.369 -0.313 -0.255 -0.213 -0.194 -0.200

0.540 -0.046 -0.116 -0.159 -0.172 -0.145 -0.313 -0.357 -0.354 -0.268 -0.237 -0.213

0.600 -0.020 -0.085 -0.122 -0.132 -0.093 -0.164 -0.215 -0.358 -0.287 -0.260 -0.229

0.700 -0.002 -0.067 -0.096 -0.098 -0.054 -0.070 -0.040 -0.127 -0.283 -0.324 -0.297

0.800 -0.024 -0.079 -0.108 -0.108 -0.063 -0.074 -0.034 -0.006 -0.078 -0.213 -0.204

0.900 -0.031 -0.083 -0.116 -0.109 -0.058 -0.035 0.010 0.030 0.010 -0.101 -0.243

0.960 -0.138 -0.221 -0.287 -0.339 -0.275 -0.171 -0.090 0.010 0.090 0.040 -0.006

0.48 0.000 1.011 1.048 1.094 1.131 1.203 1.198 1.251 1.327 1.252 1.102 1.004

0.025 -0.096 -0.178 -0.203 -0.189 -0.073 0.040 0.120 0.220 0.320 0.330 0.360

0.050 -0.206 -0.308 -0.367 -0.398 -0.315 -0.177 -0.104 0.000 0.100 0.110 0.150

0.075 -0.231 -0.339 -0.418 -0.479 -0.436 -0.295 -0.223 -0.120 -0.024 -0.025 0.010

0.100 -0.231 -0.347 -0.427 -0.501 -0.541 -0.398 -0.322 -0.229 -0.127 -0.132 -0.092

0.125 -0.237 -0.346 -0.420 -0.475 -0.614 -0.478 -0.405 -0.326 -0.215 -0.225 -0.184

0.150 -0.181 -0.274 -0.338 -0.374 -0.602 -0.496 -0.423 -0.374 -0.262 -0.274 -0.238

0.210 -0.079 -0.149 -0.199 -0.216 -0.122 -0.319 -0.270 -0.256 -0.185 -0.308 -0.283

0.260 -0.046 -0.126 -0.171 -0.199 -0.183 -0.132 -0.069 -0.105 -0.060 -0.077 -0.212

0.320 -0.070 -0.139 -0.179 -0.201 -0.252 -0.244 -0.154 -0.095 -0.056 -0.064 -0.079

0.370 -0.031 -0.103 -0.144 -0.163 -0.198 -0.232 -0.183 -0.155 -0.085 -0.056 -0.048

0.420 -0.021 -0.090 -0.129 -0.1 44 -0.122 -0.300 -0.225 -0.157 -0.074 -0.076 -0.043

0.470 -0.011 -0.086 -0.127 -0.139 -0.112 -0.273 -0.217 -0.202 -0.128 -0.102 -0.082

0.540 -0.017 -0.081 -0.118 -0.129 -0.101 -0.336 -0.280 -0.195 -0.158 -0.140 -0.108

0.600 -0.016 -0.073 -0.107 -0.112 -0.077 -0.141 -0.318 -0.246 -0.185 -0.163 -0.140

0.700 -0.005 -0.069 -0.097 -0.100 -0,058 -0.080 -0.054 -0.307 -0.218 -0.197 -0.205

0.800 -0.012 -0.069 -0.099 -0.098 -0.055 -0.074 -0.023 --0.013 -0.235 -0.235 -0.228

0.900 -0.050 -0.105 -0.130 -0. ! 27 -0.078 -0,036 0.030 0.070 0.000 -0.158 -0.214

0.960 -0.087 -0.160 -0.202 -0.216 -0.170 -0.134 -0.061 0.050 0.090 -0.085 -0.143

0.86 0.000 0.970 1.012 1.059 1.093 1.167 1.173 1.178 1.270 1.224 1.079 0.980

0.025 -0.090 -0.187 -0.239 -0.253 -0.159 -0.025 0.030 0.150 0.250 0.270 0.300

0.050 -0.186 -0.288 -0.359 -0.409 -0.388 -0.253 -0.175 -0.068 0.020 0.040 0.080

0.075 -0.207 -0.308 -0.384 -0.446 -0.451 -0.320 -0.242 -0.170 -0.084 -0.057 -0.017

0.100 -0.201 -0.308 -0.382 -0.449 -0.525 -0.393 -0.317 -0.246 -0.166 -0.168 -0.125

0.125 -0.186 -0.270 -0.339 -0.384 -0.558 -0.439 -0.365 -0.299 -0.218 -0.223 -0.182

0.150 -0.151 -0.242 -0.300 -0.331 -0.569 -0.482 -0.408 -0.350 -0.266 -0.271 -0.238

0.210 -0.043 -0.120 -0.163 -0.177 -0.119 -0.290 -0.249 -0.256 -0.180 -0.199 -0.208

0.260 -0.030 -0.102 -0.141 -0.156 -0.138 -0.082 -0.017 -0.165 -0,152 -0.191 --0.190

0.320 -0.007 -0.078 -0.122 -0.139 -0.149 -0.128 -0.072 -0.021 0.010 -0.126 -0.205

0.370 -0.026 -0.091 -0.125 -0.139 -0.138 -0.175 -0.113 -0.075 -0.008 0.010 -0.040

0.420 -0.013 -0,075 -0.110 -0.123 -0,106 -0.173 -0.117 -0.101 -0,034 -0,013 0.010

0.470 -0.005 -0.069 -0.104 -0.115 -0.091 -0.231 -0,162 -0.087 -0.067 -0.065 -0.030

0.540 -0.010 -0.071 -0.105 -0.113 -0.085 -0.235 -0.189 -0.152 -0.087 -0.077 -0.073

0.600 -0.006 -0.065 -0.098 -0.101 -0.067 -0.260 --0.202 --0.139 -0.104 -0.100 -0.078

0.700 0.000 -0.062 -0.094 -0.099 -0.063 -0.088 -0.088 -0.207 -0.143 -0.134 -0.112

0.800 -0.011 -0.063 -0.097 -0.094 -0.055 -0.066 -0.022 -0.171 -0.175 -0.156 -0.133

0.900 -0.016 -0.072 -0.094 -0.078 -0.001 0.060 0.120 0.160 0.040 -0.026 -0.023

0.960 -0.120 -0.182 -0.215 -0.212 -0.150 -0.112 -0.055 -0.012 -0.082 -0.241 -0.248

39
Table C-3. Steady-state chordwise pressure distribution data obtained in flight, lip - 45,000 ft.

Free-stream Mach number

z,ao• x/c 0.70 0.79 0.90 0.98 1.06 1.16 1.28 1.39 1.49 1.60 1.69 1.80 1.90 2.00

0.16 0.000 1.110 1.147 1.220 1.220 1.232 1.339 1.447 1.450 1.431 1.370 1.461 1.499 1.424 1.317

0.025 -0.202 -0.190 -0.103 -0.006 0.060 0.190 0.260 0.260 0.270 0.290 0.360 0.390 0.420 0.420

0.050 -0.334 -0.351 -0.300 -0.202 -0.138 --0.017 0.060 0.040 0.070 0.110 0.170 0.220 0.250 0.270

0.075 -0.360 -0.397 -0.366 -0.261 -0.215 -0.107 -0.019 -0.058 -0.029 0.000 0.060 0.110 0.140 0.180

0.100 -0.368 -0.409 -0.414 -0.304 -0.268 -0.176 -0.080 -0.157 -0.119 -0.076 -0.023 0.020 0.060 0.090

0.125 -0.346 -0.387 -0.413 -0.309 -0.258 -0.165 -0.068 -0.215 -0.182 -0.141 -0.089 -0.038 -0.005 0.020

0.150 -0.294 -0.328 -0.465 -0.347 -0.282 --0.177 -0.070 -0.171 -0.224 --0.192 -0.143 --0.093 -0.057 -0.024

0.210 --0.206 -0.238 -0.201 --0.182 -0.121 -0.012 0.090 0.100 0.040 -0.003 0.010 0.030 0.030 0.020

0.260 -0.261 -0.322 --0.282 -0.349 -0.321 -0.223 -0.104 -0.045 -0.042 -0.043 0.000 0.040 0.070 0.080

0.320 -0.220 -0.263 -0.382 -0.375 -0.414 -0.340 --0.222 -0.157 -0.117 --0,113 -0.056 -0.004 0.030 0.050

0.370 -0.180 -0.213 -0.479 .-0.422 -0.381 -0.311 -0.206 -0.168 -0.133 -0.121 -0.074 -0.020 0.010 0.040

0.420 -0.164 -0.188 -0.181 -0.370 -0.361 -0.288 -0.240 -0.219 -0.174 --.0.143 -0.097 -0.053 -0.018 0.010

0.470 -0.162 -0.188 -0.166 -0.382 -0.353 -0.292 --0.206 --0.188 -0.193 -0.176 -0.127 -0.092 -0.058 -0.024

0.540 -0.161 --0.173 -0.165 -0.242 -0.438 -0.390 -0.286 -0.243 -0.215 -0.219 -0.186 -0.152 -0.117 -0.080

0.600 -0.125 -0.132 --0.110 -0.119 --0.290 -0.401 -0.315 -0.267 -0.234 --0.225 -0.194 --0.163 -0.141 -0.111

0.700 -0.100 -0.100 --0.076 -0.054 -0.110 -0.167 -0.311 -0.325 -0.290 --0.258 -0.224 -0.192 --0.171 -0.165

0.800 -0.111 -0.110 -0.086 -0.050 -0.059 -0.052 -0.088 -0.222 -0.205 -0.233 -0.195 -4).t61 -0.147 -0.146

0.900 -0.125 -0.122 -0.081 -0.025 -0.024 0.000 0.000 -0.087 -0.236 -0.244 -0.209 -0.166 -0.142 -0.123

0.960 -0.314 -0.363 --0.286 -0.193 -0.124 -0.001 0.090 0.050 --0.016 -0.144 -0.110 -0.075 --0.062 -0.053

0.48 0.000 1.120 1.153 1.225 1.226 1.251 1.335 1.426 1.431 1.424 1.505 1.503 1.443 1.372 1.270

0.025 -0.220 -0.189 -0.079 0,020 0.090 0.210 0.310 0.340 0.360 0.380 0.450 0.510 0.510 0.530

0.050 -0.387 -0.405 -0.326 -0.210 -0.134 ..-0.006 0.080 0.120 0.140 0.170 0.230 0.280 0.320 0.320

0.075 -0.428 -0.491 -0.452 -0.331 -0.260 -0.130 -0.034 -0.013 0.010 0.050 0.100 0.160 0.190 0.210

0.100 -0.431 -0.513 -0.561 -0.436 -0.368 -0.237 -0.135 -0.119 -0.085 --0.052 0,000 0.050 0.090 0.120

0.125 -0.436 -0.493 -0.643 -0.523 -0.462 -0.340 -0.231 -0.217 .-O. 180 -0.134 -0.082 -0.027 0.000 0.030

0.150 -0.345 -0.377 -0.636 --0.528 -0.493 -0.380 -0.273 -0.273 --0.234 -0.193 -0.143 --0.094 -0.055 -0.026

0.210 -0.195 -0.218 -0.100 -0.339 -0.340 -0.280 -0.196 -0.322 -0.277 -0.231 -0.180 -0.132 -0.101 -0.068

0.260 -0.174 -0.202 -0.164 -0.134 -0.141 -0,126 -0.065 -0.081 --.0.229 -0.256 -0.206 -0.148 -0.122 -0.089

0.320 -0.203 -0.219 -0.293 -0.260 -0.197 -0.135 -0.074 -0.063 -0.086 -0.148 -0.111 -0.074 -0.059 -0.079

0.370 -0.153 -0.170 -0.320 -0.240 -0.266 -0.183 -0.095 -0.052 -0.045 -0.132 -0.100 -0.071 -0.050 -0.029

0.420 -0.132 -0.149 -O. 143 -0.307 -0.267 --O. 180 -0.082 -0.053 -0.024 -0.103 -0.074 --0.056 -0.043 -0.027

0.470 .-O.130 -0.143 -0.121 --0.270 -0.278 -0.235 --0.136 -0.102 -0.079 -0.082 -0.064 --0.037 -0.022 -0.015

0.540 -0.122 -0.137 -0.119 --0.295 -0.303 -0.246 -0.172 --0.133 -0.101 -0.097 -0.063 -0.057 -0.036 -0.018

0.600 -0.116 -0.117 -0.093 --0.106 -0.364 -0.277 -0.195 -0.169 -0.143 -0.125 -0.090 -0.062 -0.056 -0.034

0.700 -0.111 -0.105 -0.080 -0.049 -0.114 -0.322 --0.242 -0.196 -0.196 -0.177 -0.134 -0.105 --0.087 -0.083

0.800 -0.109 -0.102 --0.073 --0.035 -0.046 -0.064 -0.260 -0.241 -0.230 -0.220 -0.176 --0.136 -0.110 -0.106

0.900 -0,152 -0.145 -0.096 -0.024 0.000 0.040 -0.008 -0.168 -0.199 -0.231 -0.196 --0.145 -0,127 -0.110

0.960 -0.225 -0.232 -0.171 -0.114 -0,066 0.030 0.080 -0.066 -0.144 --0.178 -0.157 -0.124 --0.104 -0.090

0.86 0.000 1.091 l.t21 1.196 1.198 1.187 1.313 1.433 1.435 1.398 1.400 1.472 1.414 1.347 1.247

0.025 -0.218 -0.229 -0.148 -0.044 0.030 0.150 0.260 0.290 0.300 0.300 0.370 0.370 0.400 0.420

0.050 -0.348 -0.399 --0.383 -0.278 -0.194 -0.074 0.020 0.060 0.090 0.120 0.180 0.230 0.260 0.290

0.075 -0.374 -0.442 -0.447 -0.340 -0.278 -0.174 -0.091 -0.052 -0.017 0.010 0.070 0.120 0.150 0.180

0.100 -0.374 -0,439 -0.526 -0.414 -0.355 -0.254 -0.173 -0.151 -0.115 -0.078 -0.023 0.020 0.060 0.090

0.125 -0.345 -0.403 -0,573 -0.459 -0.409 --0.305 -0.218 -0.210 -0.177 -0,146 -0.094 -0.043 -0.013 0.010

0.150 -0.289 -0.324 -0.597 --0.497 --0.452 -0.351 -0.266 -0.259 -0.230 -0.199 -0.149 -0.098 -0.067 -0.042

0.210 --0.153 -0.173 --0.063 -0.331 -0.336 -0.260 -0.187 -0,211 -0.205 -0.204 -0.161 -0.113 -0.087 -0.070

0.260 -0.139 -0.156 -0.132 -0.087 -0.148 -0.221 -0.157 -0.173 -0.174 -0.165 -0.131 -0.094 --0.072 -0.051

0.320 -0.t13 -0.136 --0.181 -0.137 -0.126 -0.031 0.000 -0.131 -0.197 -0.171 -0.130 -0.094 -0.072 -0.046

0.370 -0.129 -0.142 --0,203 -0A82 -0.151 -0.107 -0.014 0.010 -0.055 -0.184 -0.139 -0.098 -0.078 -0.060

0.420 -0.106 -0,125 -0.136 -0.170 -0.183 -0.120 -0.041 -0.010 0.010 -0.111 -0.119 -0.100 -0.080 -0.056

0.470 -0.102 -0.116 -0.109 -0.231 -0.181 -0.150 -0.072 -0.061 -0.023 --0.076 -0.048 -0.022 -0.019 -0.041

0.540 -0.102 -0.114 -0.103 -O.215 -0.219 -0,196 -0.089 -0.075 -0.076 -0.054 -0.038 -0.020 -0.001 0.010

0.600 -0.097 -0.105 -0.088 -0.150 -0.237 -0,166 -0.121 -0.091 --0.073 -0.08l -0.044 -0.035 -0.016 0.000

0.700 -0.083 -0.093 -0.075 -0.042 -0.322 -0.218 -0.157 -0.123 -0.106 -0.110 -0.077 -0.048 --0.042 -0.027

0.800 -0.095 -0.097 -0.072 --0.032 -0.067 -0.251 -0.173 -0.154 -0.135 -0.139 -0.100 -0.060 -0.050 -0.047

0.900 --0.085 -0.072 0.000 0.080 0.100 0.130 0.010 -0.041 --0.033 -0.149 -0.058 0.000 0.010 0.030

0.960 -0.228 -0.224 -0.165 -0.101 -0.070 -0.040 -0.105 -0.246 -0.244 -0.330 -0.282 -0.181 --O. 180 -0.183

Cpmax 1.12 1.17 1.22 1.26 1.31 1.37 IA3 1.48 1.52 1.56 1.59 1.62 1.64 1.66

40
Table C-4. Chordwise pressure distribution data, obtained in flight, as a function of angle of sideslip, Mach = 0.52 and 0.70.

M = 0.52, Hp = 15,000 ft M = 0.70, Hp = 45,000 ft

13 8.04 3.95 0.00 -3.98 -8.00 8.02 4.01 -0.06 -3.98 -7.99

z/b" x/c

0.16

0.48

0.86

0.000 0.800 0.950 i .007 0.940 0.790

0.025 --0.564 --0.359 -0.136 0.060 0.260

0.050 -0.495 --0.363 -0.213 ---0.070 0.080

0.075 -0.446 --0.339 -0.226 -0.111 0.020

0.100 -0.433 -0,341 -0.237 -0.138 -0.022

0.125 --.0.403 -0.310 -0.218 -0.132 -0.022

0.150 -0.341 -0.260 -0.173 -0.094 0.000

0.210 -0.231 --0.173 -0.104 -0.043 0.040

0.260 -0.233 -0.186 -0.133 -0.085 -0.010

0.320 -0.181 -0,143 -0.105 -0.065 0.000

0.370 -0.137 -0. i 16 -0.083 -0.050 0.000

0.420 -0.108 -0,090 -0.067 -0.043 0.010

0.470 -0. ! 01 -0.089 -0.075 -0.045 0.000

0.540 -0.089 -0.080 -0.059 -0.036 0.000

0.600 -0.057 -0.056 -0.037 -0.016 0.020

0.700 -0.032 -0.026 -0.018 -0.003 0.030

0.800 -0.031 -0.030 -0.027 -0.019 0.000

0.900 -0.042 -0.037 -0.038 -0.033 -0.015

0.960 -0.193 -0,189 -0.174 -0.186 -0.176

0.000 0.780 0.950 1.007 0.940 0.750

0.025 -0.628 -0.395 -0.152 0,060 0.270

0,050 -0.542 -0.409 -0.250 -0.095 0.070

0.075 -0,490 -0.390 -0.272 -0.154 -0.019

0.100 -0.434 -0.361 -0.276 --0.182 -0.061

0.125 -0.395 -0.347 -0.273 -0,198 -0.095

0.150 -0.307 -0.262 -0.210 -0.148 -0,063

0.210 -0.160 -0.137 -0.098 -0.057 0.000

0.260 -0.117 -0.095 -0.066 -0.036 0.010

0.320 -0.115 -0.100 -0.078 -0.046 -0.001

0.370 --0.076 -0.067 -0.047 -0.023 0.020

0.420 -0.061 -0.054 -0.042 -0.015 0.010

0.470 -0.055 -0.046 -0.037 -0.017 0.020

0.540 -0.042 -0.039 -0.030 -0.017 0.010

0.600 -0.028 -0.033 -0.021 -0.010 0.020

0.700 -0.021 -0.027 -0.019 -0.010 0.010

0.800 -0.023 -0.027 -0.021 -0.013 0.010

0.900 -0.044 -0,052 -0.050 --0.045 --0.027

0.960 -0,094 -0.101 -0.097 -0.098 -0.093

0.000 0.830 0.950 0.970 0.900 0.710

0.025 -0.502 -0,328 -0.154 0.010 0.190

0.050 -0.416 -0.337 -0.227 -0.117 0.000

0.075 --0.362 -0,315 -0.232 -0,164 -0.062

0.100 -0.332 -0,295 -0.234 -0.179 -0.093

0.125 -0.273 -0.247 -0.199 -0.171 -0.104

0.150 -0.224 -0.207 -0.171 -0,143 -0,084

0.210 -0.116 -0.084 -0.066 -0.044 -0.008

0.260 -0.080 -0.060 -0.043 -0,027 0.000

0.320 -0.061 -0.045 -0.027 -0.013 0.010

0.370 --0.067 -0.052 -0.033 -0.021 0.000

0.420 -0.049 -0.032 -0.027 -0.012 0.010

0.470 -0.043 -0.033 -0,014 -0.012 0.010

0.540 -0.052 -0.032 -0.022 -0.011 0.000

0.600 -0.051 -0.024 -0.013 -0.005 0,010

0.700 -0.066 -0.029 -0.013 --0.011 0.000

0.800 -0.073 -0.029 -0.012 -0.010 0.000

0.900 -0.138 -0.039 -0.023 -0.053 -0.062

0.960 -0,159 -0.142 -0,119 -0.124 -0.107

0.920 1.046 1.089 1.058 0.910

-0.749 -0.487 -0.212 0.000 0.230

--0.667 -0.503 -0.323 -0.168 0.010

-0.598 -0.487 -0.342 -0.225 -0.058

-0.591 -0.472 -0.350 -0.248 -0.092

-0.566 -0.446 -0.323 -0.227 -0.086

-0.496 -0.382 -0.276 -0.186 -0.053

-0.347 -0.259 -0.190 -0.124 0.000

-0.367 -0.304 -0.231 -0.179 -0.081

-0.291 -0.237 -0.197 -0.154 -0.061

-0.230 -0.201 -0.166 -0.128 -0.053

-0.196 -0.170 -0.146 -0,119 -0.053

-0.183 -0.180 -0.152 -0,125 -0.072

-0.175 -0.161 -0.155 -0.126 -0.064

-0.141 -0.121 -0.114 -0.088 -0.045

-0.106 -0.096 -0.096 -0.074 -0.034

-0.118 -0.112 -0.105 -0.099 -0.064

-0.118 -0.121 -0.120 -0.115 -0.080

-0.318 -0.304 -0.288 -0.286 -0.273

0.930 1.061 1.096 1.053 0.880

-0.852 -0.529 -0.227 0.000 0.250

-0.775 -0.575 -0.374 -0.198 0.000

-0.677 -0.550 -0.408 -0.277 -0.098

-0.606 -0.509 -0.405 -0.306 -0.160

-0564 -0.491 -0.412 -0.333 -0.189

-0.442 -0.384 -0.325 -0.264 -0.155

-0.261 -0.235 -0.193 -0.137 -0.058

-0.218 -0.190 -0.162 -0.119 -0.042

-0.235 -0.220 -0. ! 90 -0.163 -0.088

-0.169 -0.159 -0,134 -0.105 -0.053

-0.146 -0.140 -0.120 -0.097 -0.050

-0.138 -0.124 -0.118 -0.100 -0.034

-0.130 -0.129 -0,117 -0.090 -0.053

-0.123 -0.105 -0.110 -0.088 -0.045

-0,113 -0.118 -0.100 -0.083 -0.054

-0.100 -0.105 -0,093 -0.091 -0.055

-0.135 -.-0.132 -0.140 -0.125 -0.107

-0.198 -0.209 -0.207 -0.207 -0.200

0.930 1.063 1.074 1.009 0.8 I0

-0.680 -0.442 -0.208 -0.023 0.200

-0.606 -0.474 -0.333 -0.208 -0.055

-0.512 -0.446 -0.350 -0.261 -0.143

-0.487 -0.425 -0,354 -0.292 -0.185

-0.413 -0.364 -0,316 -0.273 -0.186

-0.352 -0.319 -0.270 -0.244 -0.166

-0.222 -0.168 -0.133 -0.117 -0.051

-0.188 -0.153 -0,127 -0.105 -0.057

-0.159 -0.127 -0.102 -0.085 -0.040

-0.183 -0.146 -0.126 -0.102 -0,070

-0.161 -0.124 -0.100 -0.088 -0.047

-0.158 -0.117 -0.096 -0.072 -0.056

-0.178 -0.114 -0.096 -0.084 -0.054

-0.183 -0.112 -0.089 -0.077 -0,052

-0.197 -0.107 -0.073 -0.067 -0.040

-0.201 -0.121 -0.091 -0.084 -0.062

-0.286 -0.154 -0.083 -0.099 -0.119

-0.284 --0.264 -0,224 -0.219 -0.t92

41
Table C-5. Chordwise pressure distribution data, obtained in flight, as a function of angle of sideslip, Mach = 0.89 and 1.28.

M = 0.89, lip = 45,000 ft

13 7.96 3,99 -0.03 -3.82 -7.97

z//r x/c

0.16 0.000 1.112 1.189 1.216 1.188 1.104

0.025 -0.537 -0358 -0.111 0.090 0.300

0.050 -0.669 -0.524 -0.306 -0.140 0.050

0.075 -0.641 -0.524 -0,375 -0.237 -0.061

0.100 -0.641 -0.545 -0.415 -0.293 -0.124

0.125 -0.639 -0.552 -0.423 -0.281 -0.129

0.150 -0.712 -0.596 -0.472 -0.320 -0.104

0.48

0.86

M = 1.28, lip = 45,000 ft

3.57 1.93 -0.07 -2.02 -3.59

1.407 1.445 1.411 1.360 1.38 I

0.090 0.200 0.360 0.420 0.510

-0.104 0.000 O. 150 O. 190 0.280

-0.192 -0.077 0.040 0.070 O. 130

-0.224 -0.1 ! 4 -0.056 -0.025 0.020

-0.183 -0.098 -0.117 -0.027 0.040

-0.149 -0.083 .-0. i 60 -0.002 0.050

-0.054 0.060 O. 180 O. 120 O. 180

-0.176 -0.105 0.020 -0.055 -0.001

-0.294 -0.230 -0.063 -0.178 -0.127

-0.277 -0.231 -0.083 -0.174 -0.122

-0.250 -0.241 -0.137 -0.218 ---0.175

-0.242 -0.210 -0.139 -0.194 -0.139

-0.324 -0.280 -0.170 -0.268 -0.224

-0.353 -0.319 -0.194 -0.285 -0.239

-0.322 -0.326 -0.253 -0.260 -0.2{}4

-0.122 -0.097 -0.160 -0.094 -0.062

-0.031 -0.005 -0.177 0.000 0.010

0.020 0.060 0.030 0.090 O. 110

0.210 -0.606 -0.422 -0.186 -0.137 -0.046

0.260 -0.684 --0.393 -0.288 -0.260 --0.191

0.320 -0.565 -0.324 -0.378 -0.323 -0.265

0.370 -0.430 -0.503 -0.471 -0.286 -0.1 ! 8

0.420 -0.214 -0.195 -0.170 -0.138 -0.102

0.470 -0.225 -0.197 -0.164 -0.147 -0.120

0.540 -0.209 -0.180 -0.164 -0.14,4 -0.109

0.600 -0.164 -0.130 -0.114 -0.094 -0.059

0.700 .-0.133 -0.091 -0.076 -0.057 -0.028

0.800 -0.127 -0.098 -0.082 -0.075 -0.045

0.900 -0.111 -0.100 -0.082 -0.064 -0.043

0.960 -0.419 -0.351 -0.298 -0.280 -0.228

0.000 !.110 1.195 1.220 1.192 1,091

0.025 -0.594 -0.322 -0.091 O. 110 0.330

0.050 -0.690 -0.554 -0.341 -0.154 0.040

0.075 -0.808 -0.664 -0.468 -0.293 -0.110

0.100 -0.892 -0.756 -0.570 -0.419 -0.227

0.125 -0.950 -0.823 -0.654 -0.502 -.0.315

0.150 -0.889 -0.780 -0.645 -0.497 -0.282

0.210 -0.426 -0.206 -0.106 -0.129 -0.100

0.260 -0.198 -0.155 -0.179 -0.181 -0.133

0.320 -0.311 -0.313 -0.296 -0.248 -0.186

0.370 -0.241 -0.305 -0.304 -0.198 -0.106

0.420 -0.145 -0.145 -0.133 -0.118 -0.087

0.470 -0.153 -0.143 -0.130 -0.113 -0.091

0.540 -0.138 -0.137 -0.119 -0.107 -0.085

0.600 -0.107 -0.104 -0.099 -0.086 -0.056

0.700 -0.096 -0.089 -0.085 -0.070 -0.043

0.800 -0.088 -0.087 -0.075 -0.061 -0.033

0.900 -0.104 -0.115 -0.102 -0.075 -0.033

0.960 -0.163 -0.160 -0.184 -0.185 -0.189

0.000 1.104 1.183 1.188 1.138 1.025

0.025 -0.517 -0.357 -0.159 0.020 0.240

0.050 -0.768 -0.593 -0.398 -0.232 -0.054

0.075 -0.714 -0.603 -0.456 -0.343 -0.206

O. 1O0 -0.696 -0.633 -0.535 -0.450 -0.322

O, 125 -0.717 -0.654 -0.579 -0.498 -0.378

0.150 -0.740 -0.687 -0.585 -0.459 -0.329

0.210 -0.321 -0.097 -0.085 -0.126 -0.140

0.260 -0.125 -0.124 -0.145 -0.146 -0.143

0.320 -0.196 -0.181 -0.176 -0.153 -0.117

0.370 -0.222 -0.205 -0.184 -0.147 -0.115

0.420 -0.160 -0.152 -0.125 -0.108 -0.094

0.470 -0.147 -0.131 -0.106 -0.100 -0.090

0.540 -0.156 -0.129 -0.098 -0.091 -0.089

0.600 -0.153 -0.112 -0.089 -0.071 -0.071

0.700 -0,169 -0.107 -0.076 -0.065 -0.061

0.800 -0,177 -0.102 -0.068 -0.063 -0.059

0.900 -0.122 -0.027 -0.004 -0.018 -0.027

0.960 -0.274 -0.208 -0.172 -0.166 -0.177

1.430 1.443 1.430 1.378 1.374

O. 170 0.270 0.430 0.450 0.530

-0.049 0.040 0.210 0.210 0.290

-0.172 -0.083 0.070 0.070 0.140

-0.273 -0. ! 80 -0.027 -0.047 0.010

-0.364 -0.265 -0.121 -0.152 -0.092

-0.413 -0.308 -0.176 -0.213 -0.160

-0.256 -0.212 -0.216 -0.134 -0.073

-0.127 -0.083 -0.028 -0.026 0.020

-0.082 -0.063 -0.024 -0.064 -0.027

-0.102 -0.090 -0.012 -0.094 -0.058

-0.095 -0.076 -0.009 -0.088 -0.052

-0.145 -0,140 -0.049 -0.150 -0.110

-0.181 -0.168 -0.084 -0.168 -0.127

-0.219 -0.192 -0.113 -0.199 -0.174

-0.269 -0.242 -0.170 -0.226 -0 ! 97

-0.315 -0.289 -0.193 -0.168 -0.115

0.000 0.000 -0.124 -0.008 0.000

0.070 0.070 -0.113 0.100 0.140

1.429 1.450 1.418 1.332 1.326

O. 130 0.230 0.370 0.370 0.460

-0.077 -0.002 O. 150 O. 130 0.200

-0.184 -0.120 0.030 0.000 0.060

-0258 -0.205 -0.063 -0.104 -0.051

-0298 -0.244 -0.129 -0.177 -0.131

-0.329 -0.280 -0.189 -0.238 -0.194

-0.270 -0.192 -0.167 -0.209 -0.173

-0.258 -0.179 -0.111 -0.146 -0.072

0.000 0.000 -0.119 0.000 0.030

-0.008 -0.001 0.040 -0.042 -0.018

-0.038 -0.030 0.030 -0.082 -0.068

-0.061 -0.064 -0.018 -0.105 -0.084

-0.106 -0.088 -0.056 -0.117 -0.114

-0.126 -0.119 -0.077 -0.128 -0.099

-0.170 -0.149 -0.091 -0.149 -0.136

-0.219 -0.191 -0.116 -0.153 -0.136

-0.054 0.010 -0.068 0.060 0.080

-0.109 -0.100 -0.206 -0.100 -0.051

42
Table C-6. Chordwise pressure distribution data, obtained in flight, as a function of angle of attack.

z//r

M = 0.54, lip = 30,000 ft

-1.85 0.30 2.19 4.05

13 -0.21 -0.42 -0.27 -0.26

x/c

0.16 0.000 1.029 1,034 1.023 1.000

0,025 -0,127 -0.104 -0,104 -0.103

0.050 -0.210 -0.197 -0.189 -0.182

0.075 -0.221 --0.196 -0.197 -0.192

0.100 -0.238 -0.223 -0.214 -0.202

0,125 -0.215 -0.203 -0.189 --0,186

0,150 -0.167 -0.159 -0.145 -0.148

0.210 -0.103 -0.087 -0.081 -0.069

0.260 -0.134 -0.119 -0.112 -0.099

0.320 -0,102 -0.086 -0.084 -0.080

0.370 ---0,075 --0.067 --0,058 -0.054

0.420 --0,059 --0,053 -0,048 -0.050

0.470 -0.068 -0.054 -..0.050 -0.055

0,540 -0,062 --0,055 -0.047 -0.051

0.600 -0.033 -0,025 -0,027 -0,028

0,700 -0,011 -0.011 -0.011 -0.020

0.800 -0.024 -0.020 -0.030 -0,047

0.900 -0,028 -0,030 -0.043 -0,063

0.960 -0,136 -0,148 -0.166 -0.189

0.48 0.000 1,033 1.039 1,029 1.007

0,025 -0,141 -0.124 -0.120 -0.124

0.050 -0.252 -0.231 -0.232 -0,223

0.075 -0.271 -0.248 -0,248 --0.247

0.100 -0.276 -0.256 --0.250 -0,241

0.125 -0.276 -0.258 -0.257 -0.252

0.150 -0,208 -0.201 -0,189 -0,189

0.210 -0,095 -0,087 -0,074 -0.082

0.260 -0,071 --0,060 --0,054 -0,053

0,320 -0,095 -0,080 -0,079 -0.083

0.370 -0.047 --0.039 -0,037 -0.042

0,420 -0,034 -0,023 -0,025 -0,026

0,470 -0.034 -0.027 -0.027 -0,033

0,540 -0.026 -0.022 -0.026 -0.022

0.600 -0,016 --0.014 -0.018 -0,031

0.700 -0.015 -0.013 -0.015 -0.026

0,800 -0.017 -0.019 --0.024 --0.036

0,900 -0,046 -0,047 -0,059 -0,073

0,960 --0,093 -0,090 -0,110 -0,135

0.86 0.000 0,990 0.990 0.980 0.920

0.025 -0.151 --0.125 -0.106 -0,098

0,050 -0,240 --0.213 --0.207 -0.197

0.075 -0,255 -0,229 -0,217 -0.197

0.100 -0.250 -0,230 --0,211 -0,213

0,125 -0,222 -0.208 -0.197 -0.187

0.150 -0.182 -0.169 -0.166 -0,163

0,210 -0,071 --0.057 -0,061 --0.060

0,260 -0,053 -0,046 -0.041 -0,044

0.320 -0.025 --0,020 -0,023 -0.026

0.370 -0,047 -0.041 -0.042 -0,051

0.420 -0.026 -0,015 -0.026 -0.028

0,470 -0.025 -0.019 --0.022 -0,034

0.540 -0.020 -0.013 -0.026 -0.045

0.600 -0,016 -0.011 -0.023 -0.034

0,700 -0,010 --0.005 -0,014 -0,037

0.800 --0.017 -0.010 -0.030 -0.049

0.900 -0.022 -0.022 -0.033 --0.046

0.960 -0.120 -0.125 -0.142 -0.176

43
APPENDIX D

BOUNDARY-LAYER PROFILE DATA

Table D-I. Steady-state boundary-layer profile data obtained in flight, Hp - 15,000 ft.

Free-stream Math number

y (m.) 0.39 0.52 0.61 0.70 0.80 0.90 0.98

0.025 0.583 0,592 0.599 0.613 0.596 0.606 0.590

0.040 0.633 0.652 0,655 0.653 0.658 0,638 0,635

0.095 0.691 0.694 0.705 0.714 0.709 0.701 0.689

O. 150 0.709 0.714 0.728 0.740 0.743 0.738 0.713

O. 190 0.744 0.747 0.756 0.764 0.773 0,762 0.746

0.250 0.768 0,762 0.780 0.787 0.790 0.793 0.768

0.290 0.794 0.788 0.794 0.797 0.809 0.806 0.775

0.350 0.800 0.791 0.807 0.817 0.823 0.823 0.804

0.400 0.826 0.823 0.821 0.829 O.839 0.833 0.813

0.500 0.847 0.848 0.851 0.855 0.860 0,862 0.844

0.650 0.909 0.919 0.909 0.907 0.911 0.916 0.900

0.850 0.945 0.954 0.956 0.954 0.955 0.956 0.941

1.000 0.968 0.979 0.981 0.981 0,981 0.985 0.981

1.250 0.984 0,988 0.992 0.995 0,996 0.996 0,996

1.450 0.994 0,998 0.996 0,998 0.999 0,997 1.000

1.650 0.991 0.994 0.995 0.998 0.999 0,997 0,999

1.900 0,992 0.996 0.996 0.998 0,999 0.998 0.998

2.100 0.994 0.997 0.997 0.999 1.000 0.998 0.998

2.300 0.989 0.996 0.997 0,998 0,998 0,999 0.998

2.500 0.983 0.994 0.995 0.998 0.999 1.000 0.997

M r 0.42 0.53 0.64 0.76 0.85 0.95 0.99

1"/ 497.7 506.6 517.1 530.6 543.5 559.7 569.5

T,, 527.1 535.0 543.6 548.8 552.6 555.2 556.3

P,, 1194.3 1180.7 1168.2 1157.5 1153.5 1166.7 1205.5

Table D-2. Steady-state boundary-layer profile data obtained in flight. Hp = 30,000 fl.

Free-stream Mach number

y (in.) 0.51 0.62 0.70 0.79 0.89 1.00 1.05 1.15 1.28 1.39 1.48

0.025 0.566 0.582 0.580 0.583 0.576 0.562 0.535 0.551 O. 156 0.070 O. 188

0,040 0.625 0.618 0.645 0,635 0.635 0.617 0,593 0.588 0.285 0,276 0.306

0.095 0.682 0,677 0,696 0.691 0,687 0,664 0,649 0,631 0,375 0.409 0.450

0.150 0.711 0.706 0,726 0,730 0.712 0.694 0,676 0,654 0,441 0.500 0.504

0.190 0,742 0.739 0.751 0,754 0.746 0.722 0,705 0.684 0.523 0.599 0,599

0.250 0.758 0.762 0.772 0,773 0.771 0,750 0.726 0,706 0.568 0.656 0.663

0,290 0,767 0.774 0,790 0.789 0.792 0,757 0.738 0.718 0.603 0.695 0.703

0.350 0.787 0.796 0.811 0.807 0,810 0.778 0.766 0.745 0.659 0.738 0,757

0,400 0,808 0,813 0.823 0,816 0.820 0.787 0,777 0,759 0,683 0.764 0,780

0.500 0.835 0,841 0.852 0,848 0.849 0.818 0.822 0.796 0.761 0.813 0.841

0,650 0.893 0,900 0,908 0.901 0.902 0.885 0,876 0,859 0.851 0,886 0,901

0.850 0,942 0.948 0,950 0.949 0,950 0.929 0,940 0.915 0.914 0,926 0,934

1.000 0.961 0.977 0.980 0.977 0.971 0.970 0,973 0.957 0,955 0.955 0,963

1.250 0.979 0,990 0.993 0,994 0.991 0,993 0.995 0.988 0.989 0,981 0.987

1.450 0.993 0,994 0,999 0.999 0.997 0,999 0.999 0,999 1.000 1.000 0.992

1.650 0.989 0,993 0.997 0.997 0.997 0,999 1,000 1.000 0,998 0.989 0.994

1.900 0,994 0.996 0.998 0,998 0.998 0.999 0.999 0.999 0,995 0.986 0,997

2.100 1.000 1.000 1.000 1.000 0.998 1.000 0.999 0,999 0.991 0,982 1,000

2.300 0.989 0,994 0.997 0.997 1.000 0.998 0,998 0,997 0.979 0.981 0,999

2,500 0.979 0.994 0,997 0,999 0.999 0,998 0.997 0,997 0,977 0.982 0,998

M r 0.53 0.65 0.75 0.84 0.93 1.00 1.01 1.09 1.27 1.55 1.71

T/ 435.2 445.1 455.4 465.6 477.5 490.6 503.8 526.6 548.7 558.9 593.1

T K 486.2 494.7 500.3 504.6 508.9 515.1 521.5 526.7 528.4 526.5 523.1

Pw 647.6 638.8 634.4 629.2 635.0 663.7 711.3 743.1 645.0 606, t 603.4

44
Table D-3. Steady-state boundary-layer profile data obtained in flight, Hp _ 45,000 ft.

Free-stream Mach number

y (in.) 0.70 0.79 0.90 0.98 1.06 1.16 1.28 1.39 1.49 1.60 1.69 1.80 1.90 2.00

0,025 0.568 0.573 0.584 0.551 0.519 0.512 0.180 0,177 0.207 0.444 0.507 0.533 0.514 0.491

0,040 0.621 0.621 0.630 0.601 0,583 0.559 0.146 0,317 0.335 0.506 0,564 0.585 0.581 0.568

0,095 0.669 0,677 0.681 0.652 0.643 0.575 0,312 0.443 0.476 0,570 0.628 0.653 0.667 0.660

0.150 0.705 0.702 0.708 0.673 0,676 0.602 0.387 0.516 0.539 0.613 0,669 0.706 0.727 0.724

0.190 0.738 0.739 0.745 0.714 0.717 0.627 0.477 0,596 0.621 0.669 0.728 0,762 0.801 0.793

0.250 0.758 0.764 0.765 0.736 0,742 0,653 0,539 0,653 0,686 0.712 0.776 0.809 0,857 0,847

0,290 0,771 0.778 0,782 0.751 0,765 0.669 0,599 0,698 0.728 0.735 0.804 0.825 0.882 0.875

0,350 0.793 0.800 0,800 0.780 0.784 0,690 0.646 0.731 0.774 0,766 0,832 0.854 0.914 0.907

0,400 0.807 0.816 0.815 0,794 0,804 0.707 0.679 0.760 0,804 0,785 0,848 0,867 0.924 0,926

0.500 0.836 0,849 0.840 0.823 0.842 0,746 0.756 0,808 0,856 0.823 0,886 0.895 0.957 0.961

0.650 0.898 0.906 0.890 0.879 0,905 0.832 0.854 0.877 0.913 0.870 0.935 0.931 0,984 0,995

0,850 0,941 0,948 0,929 0.930 0,955 0.905 0.920 0,919 0.945 0.904 0.952 0,955 0.985 0.998

1,000 0.966 0,973 0,957 0,967 0.984 0.955 0.963 0.948 0.972 0.917 0,952 0,962 0,978 0,990

1.250 0.982 0,989 0,981 0,988 0,995 0.986 0,992 0.979 0.996 0.917 0,970 0,974 0.972 0,984

1.450 0.992 0,996 0.997 0.995 0.999 1.000 1,000 1.000 1.000 0.919 0.987 0.987 0.977 0.985

1.650 0.990 0.997 0.997 0.995 0.998 0.998 0.999 0.993 0.994 0.927 1.000 0.990 0.986 0_994

1.900 0.991 0.998 0,998 0,997 0.998 0.998 0.989 0.992 0.996 0.951 0.997 0.995 0.993 0,996

2.100 0.992 0,998 0.999 0.998 0.998 0.998 0.979 0.990 0.998 0.975 0.995 0.999 1,000 0.999

2.300 0.990 0.999 0.999 0.997 0.997 0,996 0,975 0.990 0.998 0.993 0.993 1.000 0.998 0.997

2.500 0,989 0.997 0.999 0,995 0.997 0.996 0.975 0.989 0,997 1.000 0.999 0,996 0.995 0,996

Mr 0.78 0.87 0.96 1.00 1.04 1.12 1.29 1.56 1.70 -- 1.91 2.00 2.09 2.17

T/ 427.2 438.2 451.2 461.7 475.4 496.4 510.5 530.9 553.3 602.7 620.5 635.3 652.9 679.9

T_, 446.7 445.5 445.8 446.6 449.3 454.6 459.7 467.6 476.1 586.7 587.5 587.3 585.8 582.5

P_,. 317.1 314.7 319.9 339.3 363.5 365.4 314.4 290.9 293.2 238.7 282.1 307.4 312.5 330.9

Table D-4. Boundary-layer profile data, obtained in flight, as a function of sideslip, Mach = 0.52 and 0.70.

M = 0.52, lip _ 15,000 ft M = 0.70, tip = 45,000 ft

13 8.04 3.95 0.00 -3.98 -8.00 8.02 4.01 -0.06 -3.98 -7.99

y (in.)

0.025 0.686 0.626 0.577 0.629 0,688

0.040 0.729 0,681 0,628 0.675 0.707

0.095 0.775 0,775 0,684 0.705 0.739

0.150 0,834 0.804 0.709 0.731 0,783

0,190 0,863 0,840 0.732 0,778 0.795

0.250 0.893 0.860 0.772 0.796 0.822

0.290 0.910 0.884 0,796 0.810 0.852

0.350 0.924 0,893 0.819 0.832 0.872

0.400 0,947 0,897 0.831 0.842 0,880

0.500 0,975 0.928 0.863 0.863 0,905

0.650 0,998 0.981 0,902 0.921 0.957

0.850 0,989 0.985 0.951 0.964 0.987

1.000 0.996 0.965 0.974 0.986 0,997

1.250 0,989 0.971 0.988 0.995 0.996

1.450 1.000 0.963 0.996 0,995 0,997

1.650 0,996 0.969 0.997 0.995 0,997

1,900 0,980 0.976 0,997 0,996 0.997

2.100 0.965 0.983 0,998 0.997 0,996

2.300 0.952 0.996 0,998 0,998 0.997

2,500 0.928 1.000 1.000 0,996 0,994

0.712 0.631 0.555 0.629 0.654

0.761 0.683 0.614 0.650 0.688

0.817 0.751 0.670 0.704 0,737

0,839 0.765 0.693 0,730 0,752

0.884 0.810 0.746 0,759 0.781

0.908 0.821 0.754 0.785 0.812

0.935 0844 0.773 0,796 0,832

0.955 0,875 0.795 0.819 0.839

0,973 0.877 0.814 0.837 0.864

0.985 0,907 0,843 0,875 0,899

0.999 0.952 0.909 0,921 0.950

1.000 0.955 0.947 0.963 0.973

0.993 0.944 0.972 0.982 0.991

0.996 0.945 0.992 0.988 0.987

0,999 0.970 0.993 0,989 0.999

0.985 0.981 0.991 0,988 0.990

0.966 0.986 0,989 0,989 0.991

0.947 0,990 0,986 0.991 0,991

0,935 0,960 0.973 0,990 0.997

0,922 0.936 0,960 0.991 0.994

45
Table D-5. Boundary-layer profile data, obtained in flight, as a function of angle of sideslip, Mach= 0.89 and 1.28.

M = 0.89, lip _ 45,000 fi M = 1.28, Hp _ 45,000 fl

13 7.96 3.99 -0.03 -3.82 -7.97 3.57 !.93 -0.07 -2.02 -3.59

y On.)

0.025 0.693 0.637 0.561 0.620 0.64,4

0.040 0.740 0.704 0.617 0.650 0.669

0.095 0.798 0.767 0.659 0.689 0.715

O. 150 0.834 0.799 0.685 0.7 i 8 0.739

0.190 0.880 0.848 0.734 0.750 0.773

0.250 0.916 0.878 0.763 0.779 0.800

0.290 0.938 0.899 0.779 0.792 0.819

0.350 0.961 0.913 0.795 0.809 0.838

0.400 0.973 0.928 0.811 0.826 0.853

0.500 0.986 0.957 0.834 0.858 0.889

0.650 0.999 0.988 0.882 0.905 0.935

0.850 1.000 0.989 0.925 0.947 0.977

1.000 0.998 0.983 0.952 0.975 0.990

1.250 0.998 0.981 0.978 0.990 0.995

1.450 0.993 0.986 0.993 1.000 0.997

1.650 0.983 0.994 0.993 0.996 0.997

1.900 0.959 0.996 0.996 0.997 0.997

2.100 0.935 0.999 0.998 0.998 0.998

2.300 0.905 0.997 1.000 0.997 0.996

2.500 0.858 0.992 0.998 0.999 0.996

0.157

0.270

0.411

0.504

-- 0.331 0.276 0.514

-- 0.414 0.339 0.553

0.159 0.497 0.413 0.598

O. 147 0.552 0.456 0.639

0.316 0.631 0.526 0.681

0.417 0.679 0.581 0.720

0.481 0.706 0.627 0.742

0.550 0.739 0.675 0.775

0.604 0.763 0.720 0.799

0.703 0.803 0.793 0.835

0.825 0.865 0.877 0.911

0.901 0.907 0.935 0.962

0.954 0.937 0.972 0.964

0.989 0.965 0.987 0.962

1.000 0.977 0.997 0.987

0.999 0.987 0.997 0.996

0.989 0.993 0.996 0.995

0.978 0.999 0.994 0.995

0.976 0.998 0.984 0.994

0.976 1.000 0.980 0.997

0.582

0.653

0.692

0.788

0.909

0.976

0.993

1.000

0.995

0.976

0.956

0.936

0.929

0.924

Table D-6. Boundary-layer profile data, obtained in flight, as a function of angle of

attack, Mach _ 0.54, Hp = 30,000 ft.

M = 0.54, lip _ 30,000 ft

ot -1.85 0.30 2.19 4.05

l_ -0.21 -0.42 --0.27 -0.26

y (in.)

0.025 0.588 0.583 0.572 0.599

0.040 0.636 0.631 0.631 0.68 I

0.095 0.687 0.682 0.708 0.736

0.150 0.722 0.702 0.746 0.778

0.190 0.746 0.752 0.773 0.805

0.250 0.773 0.764 0.793 0.833

0.290 0.780 0.793 0.833 0.837

0.350 0.803 0.809 0.846 0.865

0.400 0.818 0.817 0.855 0.882

0.500 0.849 0.836 0.881 0.912

0.650 0.904 0.903 0.954 0.952

0.850 0.951 0.958 0.989 0.953

1.000 0.980 0.974 0.995 0.944

1.250 0.994 0.991 1.000 0.940

1.450 1.000 0.995 0.999 0.940

1.650 0.994 1.000 0.984 0.946

1.900 0.997 1.000 0.982 0.964

2.100 0.999 1.000 0.980 0.982

2.300 0.992 0.996 0.962 0.994

2.500 0.999 0.991 0.960 1.000

46

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I. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT'rYPE AND DATES COVERED

December 1996 Technical Memorandum

4.TITLE AND SUBTITLE : S. FUNDING NUMBERS

F-15B/Flight Test Fixture II: A Test Bed for Flight Research

6. AUTHOR(S)

David. M Richwine

7.PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU 505-68-52

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2113

10. SPONSORING/MONrrORING

AGENCY REPORT NUMBER

NASA TM-4782

11. SUPPLEMENTARY NOTES

1211.DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 02

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

NASA Dryden Flight Research Center has developed a second-generation flight test fixture for use as a

generic test bed for aerodynamic and fluid mechanics research. The Flight Test Fixture II (FTF-II) is a lowaspect-ratio vertical fin-like shape that is mounted on the centerline of the F-15B lower fuselage. The fixture

is designed for flight research at Mach numbers to a maximum of 2.0. The FFF-II is a composite structure

with a modular configuration and removable components for functional flexibility. This report documents the

flow environment of the fixture, such as surface pressure distributions and boundary-layer profiles,

throughout a matrix of conditions within the F-15B/FTF-II flight envelope. Environmental conditions within

the fixture are presented to assist in the design and testing of future avionics and instrumentation. The intent

of this document is to serve as a user's guide and assist in the development of future flight experiments that

use the FTF-II as a test bed. Additional information enclosed in the appendices has been included to assist

with more detailed analyses, if required.

14. SUBJECTTERMS

Aerodynamic characterization, Boundary-layer pressures, F-I5B aerodynamic test bed,

F-15 directional stability, Flight test facility, Flight test fixture, Flow visualization,

Surface pressure distributions

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Unclassified Unclassified Unclassified

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53

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