NASA Contractor Report 177369 

LIMITED DISTRIBUTION 

Propulsion and :Airframe Aerodynamic 1 

I 

M.R. Mm 

P.E. HUey 

Date for generalreiease: August 1987. 

-- 

I> 

. _- - - 

/NASA-CB-177369) PGOEULSXLGY BNC AIRFRABE N87-2e551 dEROCYNABIC IBTERACTXCIS OF SGEERSCNIC 

STCL CChEIGUBATICNS, EBBSE 1 kina1 Beport 

cDonnell Aircraft Co.) 145 F Avail: NZIS Unclas - dG7/1!€ A - CS 3/02 COS7823 
NASA Contractor Report 177370 

Propulsion and Airframe Aerodynamic 

Interactions of Supersonic V/STOL 

Configurations 

Phase I: Final Report 

M.R. Mraz 

P.E. Hiley 

McDonnell Aircraft Company 

St. Louis, Missouri 

1 until August 1987. - 

Prepared for 

Ames Research Center 

under Contract NAS2-10791 

September 1985 

National Aeronautics and 

Space Administration 

AMES RESEARCH CENTER 

Moffett Field, California 94035 
TABLE OF CONTENTS 

Section Page 

1. INTRODUCTION. . . . . 1 

2. AERODYNAMIC CONFIGURATION DESCRIPTION . 0 5 

2-1 AERODYNAMIC CONCEPT DESCRIPTION .......... 2.2 SIMULATED V/STOL CONFIGURATION .......... 5 

9 

3. OVERALL TESTING APPROACH . . . . . 15 

3.1 BASIC DATA REQUIREMENTS .............. 15 

3.2 OVERALL MODEL CONCEPT ............... 18 

3.2.1 Elimination of Test Technique Bias Sources . 18 

a. Support System Selection ........ 18 

b. Metric Arrangement Selection ...... 24 

3.2.2 Reference Inlet and Nozzle Conditions ... 26 

3.3 SUMMARY OF TESTING APPROACH ............ 31 

4. DESCRIPTION OF COMMON MODEL COMPONENT 33 

4.1 TEST MODE CONVERSION HARDWARE 33 

4.2 METRIC BREAK SEALS ................ 36 

4.3 MODEL SUPPORT/TUNNEL INSTALLATION 40 

4.3.1 Ames 11-Foot Tunnel Installation ...... 40 

4.3.2 Ames 12-Foot Tunnel Installation ...... 43 

......... 

4.4 FORCE BALANCE ................... 46 

4.5 BALANCE THERMAL CONTROL 46 

4.6 INLET DESIGNS 57 

4.6.1 Baseline Inlet ................ 57 

4.6.2 Ultra Close-Coupled Inlet Definition .... 61 

4.7 NOZZLE DESIGNS .................. 69 

5. DESCRIPTION OF TEST MODE CONCEPTS 71 

5.1 SIMULATOR TEST MODE CONCEPTS 71 

5.1.1 Drive Air System .............. 71 

5.1.2 Bleed Air System .............. 74 

5.1.3 Flexibility Analysis ............ 75 

5.2 FLOW-THROUGH TEST MODE CONCEPT - . 79 

5.3 JET-EFFECTS TEST MODE CONCEPT . . 82 

iii 

pITw=EDU% PAGE BUNK NOT FWUED 
TABLE OF CONTENTS (Concluded) 

Section 

6. INSTRUMENTATION . . . . . . . . . . . . . 

6.1 MODEL INSTRUMENTATION . . . . . . . . . . . . . . . 6.2 CMAPS INSTRUMENTATION . . . . . . . . . . . . . . . 6.3 INSTRUMENTATION ROUTING . . . . . . . . . . . . . . 

7.1 METHODOLOGY. . . 7.2 SAMPLECASERESULTS . . 7.3 PLANNED ERROR ANALYSIS CASES 

8. REFERENCES.. . . . . . . . . . . . . . . . . . . . . . APPENDIX A. . . . . . . . . . . . . . . . . . . . . . . . . APPEXDIXB. e . . . APPENDIXC......................... 

Page 

88 

88 

88 

88 

96 

96 

97 

99 

10 1 

10 3 

106 

111 

iV 
LIST OF FIGURES 

Number Title Paqe 

1-1 Compact Multimission Aircraft Propulsion 

Simulator ..................... 

1-2 

2-1 

2-2 

2-3 

Study Aerodynamic Concepts ............ Study Baseline Aerodynamic Configuration ..... 

Effect of Thrust Vectoring on Trimmed Polar .... 

Study Ultra Close-Coupled (Modified) 

Configuration ................... 10 

2-4 Supersonic V/STOL Configuration Lift Plus 

Turbofan Lift/Cruise Concept ........... 12 

2-5 Comparison of Key Geometric Parameters for 

Supersonic V/STOL System and ANC Model ...... 13 

2 -6 External Geometry Comparison Between ANC Model 

and 9.62% V/STOL Concept ............. 14 

S-i Basic Data Requirements for Measurement of 

Propulsion/Airfrarne Interactions ......... 16 

17 

19 

21 

23 

25 

3 -2 

3-3 

3-4 

3-5 

3 -6 

3-7 

Comparison of Propulsion/kirframe Interactions . . 

Comparison of Overall Aerodynamic Performance ... Types of Support Systems .............. 

Aerodynamic Design of Support System ....... Metric Arrangement Concept ............ Aerodynamic Design of Reference Inlet 

Configuration ................... 29 

30 

32 

34 

37 

39 

3-8 

3-9 

4- 1 

4-2 

4-3 

4-4 

Reference Nozzle Geometry ............. Testing Approach Summary . Test Mode Conversion Concept ............ Seal Arrangement ................. Deflection Variation of Duct Metric Break ..... Planned Test Operating Conditions in NASA/Ames 

ll-Ft Unitary Tunnel ............... 41 

I . , ,.' 
LIST OF FIGURES (Continued) 

Number 

4-5 

4-6 

4-7 

4-8 

4-9 

4-10 

4-11 

4-12a 

4-12b 

4-13 

4-14 

4-15 

4-16 

4-17 

4-18 

4-19 

4-20 

4-21 

4-22 

4-23 

Title 

Model Installation in NASA/Ames 11-Ft Unitary 

Wind Tunnel .................... 

Page 

42 

Sting Length Effect on Adapter Induced Pressure 

Increments .................... 44 

Conceptual Model Installation in Ames 12-Ft 

Pressure Wind Tunnel ............... 45 

Model Loads and Balance Force/Pitching Moment 

Angle of Attack/Canard Deflection Test Envelope . . 48 

Balance Thermal Node Diagram ........... 50 

Envelope ..................... 47 

Balance Thermal Control System .......... 51 

TBleed=5O0F. 53 

Balance Axial Temperatures - Simulator Mode; 

Balance Axial Temperatures - Simulator Mode; 

TBleed"165OF 54 

Balance Axial Temperatures - Flow-Through Mode . . 55 

Balance Angular and Radial Temperatures ...... 56 

Articulating Cowl Lip Flap System for Baseline 

Inlet Configuration ................ 58 

Inlet Static Performance Improvement with Bellmouth for 8.5% Simulator Demonstration Model ... 59 

MCAIR Rotating Cowl Lip Configuration ....... 60 

Improved Inlet Performance with Rotating Cowl 

up ........................ 62 

Inlet Performance Comparison ........... 63 

Typical Normal Shock Inlet Stable Range . . . . 63 

Model Arranagement of Close-Coupled 

Configuration ................... 65 

Ultra Close-Coupled Inlet Configuration ...... 66 

Effect of Reverse Scarf Inlet on Flow Separation 

Bounds ...................... 67 

vi 
LIST OF FIGURES (Concluded) 

Number Title Paqe 

4-24 

4-25 

4-26 

5 -1 

5-2 

5 -3 

5-4 

5 -5 

5-6 

5-7 

5 -8 

5 -9 

5-10 

5-11 

6-1 

6-2 

6 -3 

I 

6-4 

6-5 

7-1 

7-2 

A-1 

A- 2 

B-1 

B- 2 

Effect of Inlet Contraction Ratio on Flow 

Separation Emunds ................. Effect of Leading-Edge Slat on Flow Separation 

Bounds ...................... G.E. ALBEN Nozzle Assembly ............ Simulator Test Mode Design ............ Turbine Drive Air System Schematic ........ Turbine Bleed Air System Schematic ........ Turbine Bleed System Pressure Loss System ..... Estimated Simulator Flexibility Envelope ..... Flow-Through Test Mode Design ........... Ejector Installation for Low Speed Flow-Through 

Testing ...................... Ejector Flow/Press=re P, eq~irenents Static 

Operation ..................... Jet-Effects Test Mode Design ........... CMAPS Nozzle Temperature (T7) Distortion ..... Effect of Jet Temperature on Afterbody Drag .... Model Instrumentation Summary ........... Locations for Model External Pressure 

Instrumentation .................. Locations for ALBEN Nozzle Pressure 

Instrumentation .................. Simulator Instrumentation Summary ......... Nozzle Instrumentation .............. Error Analysis Computer Program Structure 

Error Analysis Output for Sample Case ....... Turbine Drive Air System Components 

Turbine Bleed Air System Components 

..... 

........ Turbine Drive Air System Pressure Loss Analysis . . ........ Turbine Bleed Air System Pressure Loss Analysis . . 

68 

70 

70 

72 

73 

76 

77 

80 

81 

83 

84 

85 

87 

87 

89 

90 

92 

93 

95 

98 

100 

104 

105 

107 

10 9 

vii 
LIST OF ABBREVIATIONS, ACRONYMS AND SYMBOLS 

Abbreviations 

A 

AC 

Aa 

AEDC 

AE57 

AFAPL 

AFWAL 

ALBEN 

ANC 

ADEN 

AV9 

A/B 

A/C 

AV 

AR 

b 

B.L. 

E 

CD 

CL 

CM 

CP 

CR 

CT 

CMAPS 

CR 

D 

DLE 

DLI 

Area 

Inlet Capture Area 

Augmented Deflector Exhaust Nozzle 

Nozzle Throat Area 

Arnold Engineering Development Center 

CMAPS Mixer Area 

Air Force Aeropropulsion Laboratory 

Air Force Wright Aeronautical Laboratories 

Aerodynamic Load Balance Exhaust Nozzle 

Advanced Nozzle Concepts 

Average 

Afterburning 

Aircraft 

Aspect Ratio, $/Sw 

Bleed Venturi Area 

Wingspan 

Butt Line 

Mean Aerodynamic Chord 

Drag Coefficient, D/(q Sw) 

Lift Coefficient, L/(q SW) 

Pitching Moment Coefficient, PM/(qE S,) 

Compact Multbnission Aircraft Propulsion Simulator 

Pressure Coefficient, (P - P=)/q 

Inlet Contraction Ratio 

Wing Chord at Root 

Wing Chord at Tip or Aerodynamic Thrust Coefficient, 

FG/(q SW) 

Drag 

Direct Lift Engine 

Deck Launched Intercept 

DRAG COUNT D/(q Sw) = 0.0001 

EPR Engine Pressure 9atio 

viii 
LIST OF ABBREVIATIONS, ACRONYMS AND SYMBOLS 

Abbreviations 

E.S. 

F 

FS 

F-T 

G.E. 

IRAD 

J-E 

L 

L/C 

L/D 

M 

Max 

MCAIR 

Min 

MFR 

MRM 

Mt 

n 

NASA 

NPR 

P 

PCNLR 

P&WA 

PM 

PT 

pTO 

pT2 

Qtq 

Re 

STOT 

swt s 

SRM 

Engine Station 

Thrust in Drag Direction, FG cos( a+ 6,) 

Fuselage Station 

Flow Through 

General Electric 

Independent Research and Development 

Jet Effects 

Lift 

Lift/cruise 

Lift Over Drag Ratio 

Mach Number 

Maximum 

McDonnell Aircraft Company 

Minimum 

Inlet Mass Flaw Ratio 

Medium Range Missile 

Inlet Throat Mach Number 

Number of Data Measurements 

National Aeronautics and Space Administration 

Nozzle Pressure Ratio 

Pressure 

Percent Corrected Compressor Rotor Speed 

Pratt C Whitney Aircraft Company 

Pitching Moment 

Total Pressure 

Freestream Total Pressure 

Compressor Entrance Total Pressure 

Dynamic Pressure 

Reynolds Number 

Short Range Missile 

Theoretical Total Wing Planform Area 

Wing Planform Area 

ix 
LIST OF ABBREVIATIONS, ACRONYMS AND SYMBOLS (Continued) 

Abbreviations 

T Temperature 

TS Tunnel Station 

Tv Thrust Vectoring 

2 -D Two-Dimensional 

UPWr Unitary Plan Wind Tunnel 

V/STOL Vertical/Short Take-Off and Landing 

WAZR, 

M Compressor Corrected Airflow 

Wt2 

WL 

- 

Y 

Waterline 

Distance in Lateral Plane From Aircraft Centerline to E 

Greek Symbo 1s 

a 

8 

A 

6, 

~TATIC 

6V 

Am 

x 

U 

Angle of Attack 

Sideslip Angle 

Incremental Value 

Canard Deflection Angle 

Static Thrust Vector Angle 

Nozzle Deflection Angle 

Leading Edge Sweep Angle 

Taper Ratio, CR/* 

Standard Deviation 

Subscripts 

AMB Ambient Static 

j Jet 

SLS I Sea Level Static Installed 

T,t Total 

QD* 0 Freestream 

X 
SUMMARY 

A wind tunnel model of a supersonic V/STOL fighter configuration has been conceptually designed to measure the aerodynamic 

interaction effects which can result from geometrically closecoupled propulsion system/airframe components. The approach is 

to configure the model for tests representing two different test 

techniques. One is the conventional test technique, where 

absolute configuration aerodynamics (including inlet/airframe 

interactions) are measured in a flow-through mode, and incremental nozzle/airframe interactions are measured in a jet-effects 

mode. The other is a propulsion simulator technique, where a 

sub-scale, externally powered engine is mounted in the model, 

thus allowing measurement of inlet/airframe and nozzle/airframe 

interactions simultaneously. 

Comparison of the aerodynamic characteristics between t'ne 

two test techniques is a direct indication of the extent to which 

inlet and nozzle flowfields are coupled together. If significant 

coupling does exist, then the simulator test technique may be 

required in the future to properly measure the aerodynamic characteristics of compact fighter configurations. 

Measurement of these propulsion/airframe interaction effects 

is being carried out in a three phase program, sponsored by the 

NASA Ames Research Center, involving preliminary model design, 

detailed model design and fabrication, and high speed (up to Mach 

1.4) testing. In Phase 1, the preliminary design of the wind 

tunnel model has been completed. 

The aerodynamic configuration is a canard/wing concept 

designed for high transonic maneuverability, employing non-axisymmetric, vectorable exhaust nozzles located near the wing trailing 

edge. The external airframe components are existing hardware 

from an Air Force sponsored nozzle research program now in 

progress . 

xi 
The model has been extensively re-designed internally to 

meet NASA objectives, and a new strut support system has been 

designed. The model is designed for testing in the flow-through, 

jet-effects, and simulator modes under the philosphy to maximize 

hardware commonality. me netric arrangement in all modes i3 

identical, wherein a single internal balance is used to measure 

all external aerodynamic forces except on the nozzle Soattail. 

External pressures are used to measure the boattail aerodynamics. 

The commonality feature extends to all three metric break locations and seal bridging mechanisms. The purpose of commonality 

is to eliminate bias error sources between modes, so that comparison of aerodynamic data will represent inlet/nozzle coupling 

effects, not differences due to test technique. The metric 

arrangement is considered a relatively low risk approach, since 

only configuration external aerodynamics are measured by the 

balance (i .e., thrust is. not measured) I and the metric break seal 

arrangements are based on extensive recent experience at MCAIR. 

Considerable effort was also expended to design the model 

inlet systems for compatibility with the simulator up to 25" 

angle of attack at high speed, and 90' angle of attack at low 

speed. An active thermal control system was also designed for 

the internal balance, so that errors would not be introduced by 

the wide range of flow temperatures used to power the simulator. 

The model preliminary design is a product of extensive design and analysis by MCAIR propulsion and model design engineers. 

It was designed recognizing the main model purpose is for measurement of interaction effects with two uniquely different test 

techniques. Every effort was made in the design to ensure that 

any differences in the aerodynamics between test techniques is an 

effect of propulsion/airframe coupling, not test technique bias. 

This is not meant to infer that the model is simple, since this 

program represents the first time that two propulsion simulators 

of this type will be tested in a realistic wind tunnel model. We 

believe that this conceptual design provides an excellent basis 

foc the detailed model design and fabrication in Phase 2, and 

eventual successful testing in ?hase 3. 

xii 
1.0 INTRODUCTION 

Many of the configurations proposed for advanced supersonic 

V/STOL aircraft are very compact in nature. This results primarily from the design goal to minimize control forces and forward 

lift engine size by concentrating the major components of the aircraft near the center of gravity. Integration of the propulsion 

system with the airframe for these configurations can result in 

potentially significant aerodynamic flowfield interactions. 

These interactions may arise from geometrically close-coupled 

wing/canard/inlet arrangements along with minimum length nacelles 

and supercirculation induced by vectoring exhaust nozzles. 

The data obtained with conventional wind tunnel test techniques can be questionable in the presence of large flowfield 

interactions, since these techniques cannot achieve simultar_ssus 

aiz~latforr of ail of the flowfields involved. This can be 

achieved with the Compact Multi-Mission Aircraft Propulsion 

Simulator (WS), developed by the Air Force Aeropropulsion 

Laboratory (AFAPL), Reference 1. The Ws# Figure 1-1, is a 

miniature low bypass ratio turbofan engine powered by a high 

pressure air turbine. 

The most beneficial application of the CMAPS will obviously 

be on those aircraft that have potentially large flowfield interactions between the inlet, nozzle, and airframe. Since testing a 

CMAPS equipped model may be more complex and expensive than testing a conventional model (flow-through and jet-effects) , the need 

to identify the types of configurations which require CMAPS 

evaluation is critical An aerodynamically "close-coupled" 

V/STOL configuration represents an effective means of evaluating 

the requirement for simultaneous inlet and exhaust nozzle flow 

simulation, and thus the potential need for the CMAPS testing 

technique. 

1 
I I 1 

c 

x 

0 

Q 

c 

0 

2 
Based on the foregoing considerations, a three phase NASA 

program of 29 months planned duration has been initiated to 

measure airframe/propulsion system interactions on close-coupled 

supersonic V/STOL configurations using the propulsion simulator 

and conventional model techniques. A secondary objective is to 

begin development of installation and test techniques for propulsion simulator equipped wind tunnel models. 

The approach to accomplish these objectives is to design, 

fabricate, and test two model configurations characterized by a 

close-coupled and a closer coupled airframe/propulsion arrangement - each in simulator and conventional test modes. Key characteristics of the test configurations are shown in Figure 1-2. 

The basic model (close-coupled) of the baseline configuration 

(i.e. external airframe components) will be provided by the Air 

Force. It was developed under prime contract to the Air Force 

Wright Aeronautical Laboratories (MAL) by MCAIR in the Advanced 

Nozzle Concepts (ANC) program, Reference 2. The very closecoupled model is a modification of the basic configuration with 

shortened inlet length (canard removed). 

Preliminary (conceptual) design of the test model was completed in Phase 1 during the first seven months of the program. 

The essential products of Phase 1 are the design layouts of the 

model in the different test modes, and the supporting analyses. 

The layouts have been transmitted per Reference (3). The supporting analyses and other necessary documentation are the subject of 

this report. 

The remainder of the report includes description of the 

aircraft test configurations; overall testing approach: model 

common components; test mode concepts; instrumentation; and error 

analysis. 

3 
a 

c 

0 

v) 

c 

a 

W 

0 

z 

0 

0 

a 

> 

P 

c 

v) 

a 

4 
I 

2.0 AERODYNAMIC CONFIGURATION DESCRIPTION 

The baseline aerodynamic configuration for the study was 

selected in proposal studies (Reference 4) to be consistent with 

several VASA program requirements, as follows: 

( 1) Twin engine propulsion system with non-axisymmetric and 

deflectable exhaust nozzles. 

(2 Aerodynamic configuration and propulsion system integrated to maximize supercirculation lift induced by 

thrust vectoring. 

(3 ) Configurations with high probability of demonstrating 

aerodynamic coupling of inlet and exhaust nozzle flowfields, including a baseline close-coupled configuration and one even more close-coupled in nature. 

(4) Realistic design of a supersonic V/STOL fighter/attack 

aircraft utilizing a wing/canard configuration. 

In addition, other desirable requirements included engine cycle 

compatibility with the CMAPS and commonality with other programs 

to minimize cost. 

The aerodynamic concept selection involved (1) choosing an 

aerodynamic concept to meet the configuration requirements above 

(Items 1, 2 and 3) and (2) conducting sizing studies of various 

propulsion systems in the aerodynamic concept to assess the 

realism for meeting V/STOL fighter/attack requirements (Item (4) 

above), concluding in selection of a specific V/STOL system for 

test simulation. 

2.1 AERODYNAMIC CONCEPT DESCRIPTION - The aerodynamic baseline 

configuration selected for the study was derived frm a concept 

developed for the Air Force sponsored ANC program, Reference 2. 

5 
The selected configuration, Figure 2-1, is characterized by 

widespaced, podded nacelles, a canard/wing arrangement, twin vectorable 2-D nozzles at the wing trailing edge, and close-coupled 

airframe (canard/wing) and propulsion system (inlet/nozzle) 

components. The nozzle concept is a General Electric Aerodynamic 

Load Balanced Exhaust Nozzle (&BEN), with fully continuous 

vectoring capability up to 230' 

A major design objective for this aerodynamic concept was to 

provide a combination of the movable canard and vectorable 

nozzles to achieve maneuvering drag reductions. This was accomplished by tailoring the canard/wing arrangement to achieve a 

maximum unstable static margin of 15372 at subsonic speeds. With 

this relaxed stability level, the vehicle pitching moment characteristics are such that a nose-up moment is produced with the 

canard deflection for minhum maneuvering drag. Positive nozzle 

deflection (nozzle deflected downward) is then used for trim, 

producing a beneficial lift/drag increment from supercirculation. 

The estimated thrust vectoring payoff in drag is illustrated by 

canparing trim& polar estimates with and without vectoring , 

Figure 2-2. At a typical maneuvering lift coefficient of 0.8, 

the drag reduction is about 100 counts, or 10% of aircraft drag 

(thrust loss due to vectoring included). 

Maximizing the degree of coupling between the airframe and 

propulsion system components was a strong consideration in selection of the baseline configuration for the program. Of course, 

it is possible to only qualitatively assess the level of coupling, since the wind tunnel data are yet to be obtained. There 

is little question that the inlet/canard interactions will be 

significant, since the canard is located at the side. of the 

inlet. Similarly, the nozzle/wing interactions should be maximized due to the location of the vectoring nozzle at the wing 

trailing edge. 

6 
7 
cL TOT 

1.0 

0.8 6 G SUSTAINED 

TURN 30,OOO AT FT 

0.6 I/ 

0.4 // 

e- 

/- 

I /4L- KD a 0.0100 

I 

I 

DRAG COEFFICIENT - CD TOT 

OPlM 

FIGURE 2.2 

EFFECT OF THRUST VECTORING ON TRIMMED POLAR 

ANC Estimate Air-to-Air Vehicle 15% Unstable 

Mach 0.9 Max AIB (C~=0.14) 

8 
The level of geometric coupling fore to aft is dictated by 

the nacelle length. In the baseline configuration, the nacelle 

length was established by the size requirements of the closecoupled canard and wing. The related length of the inlet duct is 

equivalent to approximately four engine face diameters. The 

I 

I resulting nacelle length is 54% of the overall aircraft length. 

Comparing to the F-15, which is generally regarded as a "farcoupled" installation, the inlet duct length is about seven 

engine face diameters, and the nacelle length is approximately 

65% of overall aircraft length. Since both the study baseline 

I configuration and the F-15 are nearly equal in overall length, 

the nacelle of the close-coupled baseline is about 20% shorter 

I than the F-15 

: 

I To achieve a more closely coupled configuration, the inlet 

I 

duct was shortened to about one engine face diameter in length. 

This ultra close-coupled configurstion is shcq6:l: in ~igr;re 2-3. 

The nacelle length is about 30% shorter than the baseline. 

Considering propulsion system requirements alone, the decreased 

other change to achieve the modified configuration is to remove 

the canard. The shorter nacelle length precludes a nacellemounted canard, unless the wings are changed. 

I 

I duct length is practical since the duct is straight. The only I 

I 

Testing of both the selected baseline and modified configurations should provide an excellent approach for achieving the high 

level of propulsion/airframe coupling considered necessary to 

fulfill program objectives. 

2.2 SIMULATED V/STOL CONFIGURATION - The viability of the selected baseline aerodynamic configuration as a supersonic V/STOL 

fighter was assessed in proposal studies through sizing analyses 

on a deck launched intercept (DLI) mission. Three different 

V/STOL propulsion systems were evaluated as candidates. One was 

selected as a study baseline to allow developnent of the aerodynamic performance at inlet/nozzle flow conditions which are 

9 

consistent with a representative V/STOL system. Selection 

criteria were twofold; (1) compatibility of the required model 

scale for the V/STOL candidate with model hardware available from 

the ANC program, and (2) capability of the CMAPS to match the 

operating characteristics of the candidate lift/cruise engine. 

Based on the sizing studies and other noted considerations, 

a lift plus turbofan lift/cruise configuration, Figure 2-4, was 

selected as the study V/STOL concept for test simulation. The 

scale of a CMAPS-equipped wind tunnel model of this V/STOL configuration would be set by matching the maximum airflow of the CMAPS 

to the maximum engine airflow (normally the sea level static 

condition). Specifically, the model scale factor is determined 

as follows: 

Max CMAeS Compressor Corrected Airflow 

Max Engine Corrected Airflow Scale Factor =d 

Application of this relationship to the turbofan lift/cruise concept results in a model scale of 9.62%. 

A comparison of the key geometric parameters of the scaled 

V/STOL configuration with the existing ANC model is provided in 

Figure 2-5. The model airframe parameters (i.e. wing, canard, 

fuselage, nacelle) are somewhat larger than would be required for 

the properly scaled V/STOL concept, as is the inlet capture area. 

The nozzle throat areas are essentially correct. It is planned 

to modify the ANC model inlet for the NASA program to reduce the 

capture area to the properly scaled value (5.46 id) . 

An external geometry comparison of the modified (smaller 

inlet area) model and the lift + lift/cruise V/STOL configuration at 9.62% scale, Figure 2-6, shows that the ANC model is an 

excellent representation of the V/STOL concept. The propulsion 

system components are properly scaled in the model, as well as 

the relative location of the canard/inlet and wing/nozzle. 

11 
I' 

I 

I 

0 

1- 

I 

v) 

3 

a 

L 

12 
OVERALL FUSELAGE LENGTH 

NACELLE LENGTH 

WING SPAN 

~VIIUU nnccr 

CANARD AREA 

INLET CAPTURE AREA 

NACELLE MAXIMUM AREA 

NOZZLE THROAT AREA (DRY) 

NOZZLE THROAT AREA (AB) 

..,..I L ."C. 

--------cSUPERSONIC V/STOL 

AT 9.6296 SCALE 

5.63 FT 

2.59 FT 

3.81 FT 

4.15 FT2 

L+ L/C TURBOFAN 

0.79 FT2 -- ----- 

5.27 IN.2 

16.7 IN? 

289 IN.* 

5.08 IN.' 

__ - - .- _- 

ANC 

MODEL 

5.88 FT 

2.83 FT 

4.28 FT 

5.22 FTZ 

AIR-TO-AIR 

1.02 FT~ ----- 

5.67 IN.'" 

18 IN.' 

2.83 IN.2 

5.15 IN.2 

*Modified to 5.46 in.2 for NASA Program aPi-2a 

FIG 

COMPARISON OF KEY GEOMETI 

V/STOL SYSTEF 

JRE 2-5 

IIC PARAMETERS FOR SUPERSONIC 

I AND ANC MODEL 

13 
9.62% V/STOL CONCEPT) 

FIGURE 2.6 

EXTERNAL GEOMETRY COMPARISON BETWEEN ANC MODEL 

AND 9.62% V/STOL CONCEPT 

14 
3.0 OVERALL TESTING APPROACH 

The overali testing approach and associated data requirements were established early in the program to ensure a proper 

groundwork for fulfilling the major program objective. Basically, the major objective is to measure propulsion/airframe 

interactions using the propulsion simulator and conventional 

model techniques. 

3.1 BASIC DATA REQUIREMENTS - The basic data requirements are 

set by the need to measure the inlet/airframe and nozzle/airframe 

interactions, and to assess the extent to which these interactions are coupled together. Here, the term basic data require- 

- ments refers to the "end product" data by which the interactions 

are quantified. If inlet/nozzle flow coupling exists, the inlet/ 

airframe Ln,terzctiofis measured by the conventional f low-through 

model with a "constant nozzle condition" will not agree with the 

interactions measured with the simulator and a "variable nozzle 

condition." Likewise, the nozzle/airframe interactions measured 

by the conventional jet-effects model with a "constant inlet 

condition" will not agree with the interactions measured with the 

simulator and a "variable inlet condition." The basic data 

requirements for the two test techniques, in terms of total 

vehicle aerodynamic coefficients, are therefore as illustrated in 

Figure 3-1. If the level of disagreement between conventional 

and simulator equipped models (left and right hand sides of 

Figure 3-11 is significant, simulator testing may be required. 

Utilizing the basic set of data (Figure 3-11, the two types 

of testing can be compared either in terms of propulsion/airframe 

interactions or overall aerodynamic performance. The interaction comparisons are between either of the conv'entional modes 

and the simulator mode, and are quantified as illustrated in 

Figure 3-2. 

15 
Represents MFR/canard 

interactions. inletlnozzle 

coupling effects not 

meaund. INLETIAIRFRAME INTERACTIONS 

CONVENTIONAL FLOW-THROUGH MODE 

(CONSTANT NOZZLE CONDlTlONSl 

MFR 

NTERACTION 

'ZERO 4/ I 

CONDITION 

SIMULATOR MODE 

(VARIABLE NOZZLE CONDITIONS) 

I 

I 

I 

I I 

I MFR 

Represents NPR/SV/power I 

I setting interactions. inlet/ 

nozzle coupling effects 

not measured. NOZTLEIAIRFRAME INTERACTlONS 

Represents MFRlconard .....:::.:. interactions. inlet/nozzle 

coupling effect measured. 

SIMULATOR MODE CONVENTIONAL JET-EFFECTS MODE I 

(CONSTANT INLET CONDITIONS) I 

m 

(VARIABLE INLET CONDITIONS) 

I 

I 

I 

I 

I 

I 

1 

I 

1 

C~(CL.C~) 

I 

I 

I 

NPR NPR 

Note:, ,Ship. of arbitrary condition 

FIGURE 3-1 

BASIC DATA REQUIREMENTS FOR MEASUREMENT OF 

PROPULSlONlAlRFRAME INTERACTIONS 

DRY AND AIB. EACH 

WITH VARIABLE INLET 

CONDITIONS (8, MFRl 

I 

I 

I 

I 

16 
31 

a 

I 

LL 

88 

I 

v) 

Z 

2 

LL 

.w 

a 

f 

t 8: 

Qd 

0 

- Q 

a 

b. 

2 

17 
The full impact of any differences in the interactions is 

probably most visible frun overall aerodynamic performance 

comparisons. This involves combining the flow-through and 

jet-effects data, using appropriate inlet and nozzle "reference 

conditions", for comparison to the simulator data. Here, comparison of trimmed drag 2olars is of most interest, as illustrated 

in Figure 3-3. 

I It should be noted that all test mode data comparisons are 

I Figures 3-1 through 3-3. Other comparisons will also be of 

I moment increments, external pressure distributions, and pitching 

not expected to take the relatively simple form illustrated in 

interest, such as those between vectoring induced lift, drag, and 

moment characteristics. In the foregoing discussion, the fundamental approach for comparisons was presented: the actual data 

elements will be dictated by the observed data levels and trends. 

3.2 OVERALL MODEL CONCEPT - Valid and meaningful measurement and 

canparison of the interactions and overall performance implies at 

least two major requirements for the mbdel approach: (1) test 

technique biases must not enter into the comparisons and 

(2) appropriate inlet and nozzle reference conditions must be 

established. 

3.2.1 Elimination of Test Technique Bias Sources - There 

are two major possible sources of bias errors. These are due to 

support system and metric arrangement differences between test 

modes. The obvious solution to eliminate these bias sources is 

for the support system and metric break locations to be common in 

all test modes. This is the model approach to be used in the 

program. 

a. Support System Selection - Selection of a common support 

system concept for all test modes is dictated by several requirements. Of these, three predominate: 

18 
a 

a 

a 

e 

d 

3 

a 

W 

A 

z 

c 

> 

z 

a 

P 

5 

8 

I 

hI a 

0 

a 

z 

0 

4. 

4 

19 
Adequate volume to contain drive, bleed, and instrumen- . 

tation lines for two CMAPS units, plus strength to 

sustain loads at angles of attack up to 25' at high q 

levels. 

Minimum interference, particularly in the nacelle and 

wing regions, where aerodynamic interactions are of 

most interest. This must take into consideration the 

requirement for future low speed testing up to 90' 

angle of attack. 

Feasibility for adaptation to other types of configurations , 

There were four fundamentally different types of support 

systems considered, as illustrated in Figure 3-4. They were 

(1) wing tip, (2) upper fuselage mounted strut, (3) lower fuselage mounted strut, and (4) rear entry sting. 

Wing tip mounting was not considered feasible because the 

high volume requirements for CMAPS air would have distorted the 

wing considerably all along the span. This would conflict with 

proper measurement of interactions involving wing supercircula- 

- 

tion effects due to vectoring. Further, flexibility to adapt a 

wing tip support to other configurations makes this approach 

unattractive. 

The upper fuselage mounted strut is attractive for high 

angle of attack testing. However, for the selected test configuration, which has the inner wing attached low on the fuselage, 

CMAPS air routing from the upper fuselage to the low inner wing 

was virtually impossible without excessive fuselage distortion 

and model complication. 

20 
WING-TIP SUPPORT 

UPPER FUSELAGE 

MOUNTED STRUT 

LOWER FUSELAGE 

MOUNTED STRUT 

REAR ENTRY STING 

FlGURE'3-4 

TYPES OF SUPPORT SYSTEMS 

21 
The lower fuselage mounted strut was given strong consideration. In fact, it is the'shpport system being used for tests of 

this model in the ANC program. It is compatible with the fuselage/wing arrangement of the test configuration, and does not 

interfere with the vertical tail. The major drawbacks are that 

the fuselage interference effects at high ancjle of attack may 

become excessive. 

A direct rear entry sting is the type of support system used 

for most force and moment testing on conventional flow-through 

models. It is probably the best system for high angle of attack 

testing: and it is generally regarded as the lowest interference 

support system, although very often the aft fuselage must be distorted for sting entry. In fact, this is the major disadvantage 

of the direct rear entry sting on the selected test configuration. Excessive aft-fuselage distortion would be required to 

accommodate the support of the model plus volume for air lines 

and instrumentation. 

The selected support system is a compromise between the 

lower fuselage mounted strut and the direct rear entry sting. 

Details of the aerodynamic design are shown in Figure 3-5. 

Adequate volume and strength are provided in the same maximum 

thickness (2.25" near the model) that is required for the lower 

fuselage mounted strut. However, elimination of the vertical 

"blade" portion of the strut acts to reduce interference at high 

angles of attack. The major disadvantage compared to the other 

candidates is probable increased interference at low angles of 

attack. However, the interference effects should be confined 

largely to 'the lower aft fuselage area, which is removed from the 

wide-spaced nacelles and outer wing regions where the interaction 

effects will predominate. 

22 
- t -- - -1 

V 

I' 

Q 

f 

Y (I) 

23 
b. Metric Arrangement Selection - The main requirements of 

metric arrangement selection were for measurement of external 

aerodynamic interactions on the complete configuration, and nonmetric mounting of the simulator units and nozzles (i.e. no exit 

gross thrust on the metric portion of the model). These requireaents dictated that the airframe "skin" components be metric in 

all test modes. A partially metric Jet-effects mode, which is 

common practice, could have been a candidate. However, if aftbody flowfield changes (e.g. due to vectoring) reached the nonmetric forebody and were not measured, then the comparisons 

between simulator and conventional modes would reflect test 

technique bias, not true interaction effects. 

The selected metric arrangement concept is characterized by 

a single balance system with three metric break locations common 

to all test modes. The concept is shown in Figure 3-6. The 

external aerodynamic forces are measured by a single internal 

balance attached to the aircraft skin, except for the nozzle 

boattail, and by 52 pressure taps located on the nozzle boattail 

surface. The CMAPS and nozzle (external and internal surfaces) 

are therefore non-metric in the simulator mode, the jet-effects 

internal ducting and nozzle are non-metric in the jet-effects 

mode, and the internal duct aft of the simulator face station and 

choke are non-metric in the flow-through mode. 

There were two major reasons for choosing the single balance 

system. First and most important, this approach fully satisfied 

the program objective to measure external aerodynamic interactions. Second, compared to a multiple balance system, it is 

the most simple, cost effective, and lowest risk approach. This 

program represents the first time that - two CMAPS units will be 

tested in a compact wind tunnel model. Internal model volume is 

at a premium; packaging too much hardware inside the model could 

result in an unreasonable amount of model development time and 

high risk. 

i 

24 
Y 

w 

a 

a 

m 

Y 

a 

'\ 

\ \, 

\ 

\ 

Y 

a 

a 

m 

Y 

a 

I 

Io 

3 

w 

G 

a 

I 

i 

i 

i 

I 

I 

I 

I I 

I 

J 

L 

25 
26 

In the proposed metric arrangement (Figure 3-6), there are 

three' metric breaks and three associated seals common to all 

testing modes. The three seal locations are at the simulator 

compressor face station, (duct metric Sreak) , at the beginning of 

the nozzle boattail (aft metric break), and between the strut and 

lower fuselage (strut metric break). The strut and aft metric 

break concepts are not new, and have been successfully employed 

in recent nozzle research investigations (e.g. References 2 and 

5). The duct metric break is somewhat unconventional, since 

usual practice for conventional flow-through testing is a completely metric internal duct. However, the only practical 

location for the duct metric break in the simulator mode is at 

the simulator inlet. Here, a seal tare and a duct exit momentum 

tare must be determined. The duct metric break for the two conventional modes was also chosen to be at this same location. The 

idea is to eliminate possible bias errors due to determination of 

the tares at two different locations. Further details concerning 

the metric break configurations are presented in Section 4.0. 

3.2.2 Reference Inlet and Nozzle Conditions - Specified 

reference conditions are required primarily for overall performance determination using data from the two conventional test 

modes. In common practice, the flow-through model, tested with 

various size chokes at the duct exit and variations in the aerodynamic control surfaces, provides the basic aerodynamic polars, 

trim characteristics, and drag variations with mass flow ratio 

(MFR). Increments from the jet-effects model (often partially 

metric) are applied to the flow-through data to account for 

operating nozzle geometry and nozzle pressure ratio (NPR). In 

order to apply the increments properly, a common nozzle condition 

must be tested on the two conventional models. The comon nozzle 

condition is usually a flow-through model choke tested at the low 

NPR corresponding to a flowing inlet condition. 
One problem is often neglected in this type of conventional 

testing. The chokes used to obtain the effects of inlet MFR on 

the flow-through model also produce a variable plume shape and 

possible variation in the flow over the nozzle Soattail with MFR. 

If the model is fully metric, these plume effects can be misinterpreted as due to YFR variations and can cause errors in the 

predicted full scale performance at some i4FR conditions. 

Based on the foregoing considerations of bias sources and 

reference conditions, a testing approach concept was selected 

where "multipurpose" reference conditions are established which 

are used not only for measurement of overall performance, but for 

elimination of bias errors as well. 

Selection of the proper reference conditions presented some 

unique possibilities. The concept of a single configuration 

common to all three test modes became attractive, part-icularly 

since the same metric arrangement is used in all modes. The need 

for an inlet/nozzle configuration common to the two conventional 

test modes is readily understandable, since there must be a 

method for combining the two sets of data into overall performance. In the selected testing approach, this same configuration 

is also tested in the simulator mode. This provides a "check" 

configuration to eliminate bias from model build-up. After the 

initial build-up in the tunnel of each mode, the common reference 

configuration is tested in a series of check runs over the Mach 

number and angle of attack range. These check runs are then 

compared with runs of the common configuration on the other 

modes. If the check run data does not agree (within some established tolerances), reasons for the discrepancies must be found 

and corrected. 

There are not many choices for inlet and nozzle reference 

configurations which can be common to all three test modes. The 

inlet must be non-flowing (e.g. faired) , since inlet flow cannot 

be simulated in the jet-effects node. And since the inlet must 

27 
be non-flowing, this locks in a non-flowing reference nozzle. 

This is because there can be no simulatidn of nozzle flow in the 

flow-through mode if the reference inlet is also non-flowing. 

The choice of the geometry for the reference non-flowing 

inlet was Setween an internally blocked or externally faire6 

inlet. The blocked inlet would be high in drag, simulating a 

zero mass flow ratio condition. Based on past experience, the 

faired inlet has a more representative drag level. This is 

considered advantageous in a thrust/drag accounting system, 

otherwise, increments fran the reference performance basis are 

large, resulting in increased uncertainty at the normal operating 

drag level. Based on these considerations, the faired inlet was 

chosen as the non-flowing reference inlet condition. The 

selected fairing shapes for the baseline and ultra close-coupled 

configurations are shown in c'igure 3-7. The fairing shapes 

represent simple forward extensions of the local cowl moldline. 

There are also several factors to consider in choosing the 

exact configuration for the non-flowing reference nozzle. If the 

non-flowing condition were the only requirement, almost any 

nozzle geometry could be used. It is best, however, to make ths 

reference nozzle compatible with the nozzle requirements for the 

other testing modes. The ALBEN is not a good choice, because the 

flow-through We must hav; an easily variable throat size to 

achieve MFR variations. Further, as previously discussed, it 

would be best to remove the effect of choke size on the external 

aerodynamics so that only the effects of inlet variations will be 

measured in the flow-through mode. This is possible if the 

nozzle is extended far downstream of the normal vehicle exit 

plane, with variable chokes at the exit. I The nozzle extension at 

a no-flow condition is then used on the other test modes to serve 

as the reference nozzle geometry. This was the selected approach 

for the reference nozzle condition. The aerodynamic shape of the 

extensions are shdwn in Figure 3-8. 

28 

&++- tFIGURE 38 

REFERENCE NOZZLE GEOMETRY 

30 
The reference nozzle extensions are 'non-metric. The outer 

shape conforms to the ALBEN up to the lower flap hinge point, 

thus making the absolute drag level in the flow-through mode 

nearly representative of the actual full scale configuration. 

The circular internal flow path was chosen for ease of fabrication. Chokes of various sizes are placed at the exit of the 

extensions for flow-through testing. The extensions have the 

further advantage that they may be attached easily to an ejector 

system for low speed flowthrough testing, thus making it 

possible to achieve the proper inlet mass flow ratio. Further 

discussion of the preliminary design for the nozzle extensions 

and ejector system is provided in Section 5.2. 

3.3 SUMMARY OF TESTING APPROACH - The overall testing approach 

and data utilization features three key elements; (1) check runs 

on the common configuration, (2) comparisons of inlet/airframe 

znZ ~~zzIe/a~rfrziiiieriteractions between test modes, and (3 1 comparisons of overall performance in terms of trimed drag polars. 

These elements are illustrated in Figure 3-9. 

The check runs are made on the common reference configuration in each testing mode. Assuming no correctable model buildup biases are revealed by these tests, the results provide an 

indication of the differences due to rand- errors between the 

test modes (although without simulated propulsion flows). The 

polars for the common configuration also provide the reference 

basis for the overall performance build-up. Data from the 

"operating" conventional and simulator modes, with actual inlet 

and nozzle flows and simulated airframe variations (6=, 6~1, are 

combined with the reference configuration data to give the overall performance comparisons. The operating modes also provide 

the data for direct comparison of inlet/airframe and nozzle/ 

airframe interactions. 

31 
b 

1 z 

f 

*; Y 

r 

L 

I 

I I1 

L 

32 
I I’ 

4.0 DESCRIPTION OF COMMON..MODEL COMPONENTS 

Achievement of hardware commonality in the peliminary nodel 

design received considerable emphasis in Phase 1. The key comon 

nodel leatures, other than the fuselage/wing/canard/tail components are: (1) test mode conversion hardware, (2) metric break 

seals, (3) model support/tunnel installation, (4) force balance, 

(5) balance thermal control, (6) inlets, and (7) nozzles. 

4.1 TEST MODE CONVERSION HARDWARE - A key to implementing commonality was in design of the hardware to convert frcun the internal 

“propulsion package” of one mode to another. The test mode 

conversion concept is shown in Figure 4-1. The only components 

of the model common are the ws, flow-through duct, and 

jet-ef fects air supply plenum. “hen converting to flow-through, 

the CMAPS units are replaced with z sactisn of straight duct; 

when converting to jet-effects, the CMAPS units are replaced with 

the high pressure jet-effects plenum. All nacelle hardware forward of FS 36.33 is common, including the duct metric break seal, 

instrumentation ring, and inlet. All nacelle hardware aft of FS 

45.08 is also common, including the duct bellows, transition 

duct, and nozzles. The duct bellows is needed primarily for the 

simulator mode-acting to prevent vectored nozzle induced bending 

moments from loading the simulator. 

The real key to the test mode conversion concept is the common support hardware. This common hardware, as seen in the planview portion of Figure 4-18 ensures consistent positioning of the 

internal non-metric and metric model components. The supports 

consist of the forward side mount, bleed line manifold, and the 

nozzle support. The forward side mount and bleed line/manifold 

provide consistent alignment of the duct seal, which should 

eliminate bias in the tare correction for this seal. The forward 

side mount removes torque loads and is flexible in the axial 

direction to allow for thermal growth. Consistent sositioning at 

33 
I 

.I 
I' 

I I 

f 

1 

4 

i 

H' 

r 1 

4 

\ '\ 

1 

I 

I i. 

rti 

I 1 

I 

I 

I 

i 

1 

! 

i 

I 

35 
the aft metric break is provided by the common nozzle support and 

transition duct. This should eliminate differences in the small 

step at the ar't inecric break caused by balance deflection at 

angle of attack. A difference in step height at this metric 

break between test modes could cause a drag bias. The nozzle 

support also transmits the nozzle thrust and moment loads into 

the non-metric balance housing. 

4.2 METRIC BREAK SEALS - In t5is model concept, the airframe 

components are metric and the CMAPS units and nozzles are mounted 

non-metrically to the strut. Accordingly, there are three 

distinct bridging locations between metric and non-metric hardware. These are at the (1) strut and lower fuselage (strut 

metric break), (2) beginning of the nozzle boattail (aft metric 

break), and (3) simulator compressor face station (duct metric 

break). Details of the metric break design, including the seals, 

are shown in Figure 4-2. 

A thin flexible teflon seal will be used at the strut metric 

break. The seal is positive in that no flow enters the internal 

cavity through this juncture. This seal has a negligible installation effect, as evidenced in recent static loadings at AEDC 

prior to test of the model under the ANC program. Loadings in 

axial force, normal force, and pitching moment were made with and 

without the seal installed. There were no measurable seal 

effects, either to reduce balance output sensitivity or increase 

random error. 

The aft metric break is bridged by a floating teflon seal. 

This is not a positive seal, but simply a restrictor intended to 

stabilize pressure inside the model cavity. At the start of the 

nozzle boattail, where the seal is located, a sizeable pressure 

differential can exist between top and bottom at angle of attack. 

Therefore, to avoid flow within the cavity, some seal or 

restrictor is required. The teflon seal is attached in a groove 

to the metric aftbody and is backed by foam in the groove (see 

36 
FORWARD DUCT SEAL 

Jl 

W 

STRUT SEAL 

FIGURE 4.2 

SEAL ARRANGEMENT 

AFT METRIC SEAL 

37 
Figure 4-21. This has the effect of "spring loading" the seal 

slightly against the flat surfaces of the non-metric nozzle. The 

non-metric, downstream side of the seal must ride on a flat 

surface (i.e., as opposed to another groove) because the nonmetric nozzle deflects relative to the metric aftbody under 

vectoring loads. This seal design has also Seen shown to have a 

negligible installation effect based on the recent static check 

loadings of the ANC model. 

The duct metric break is bridged by a Butyl rubber seal. 

This positive seal arrangement is the same type used on the 8.5% 

Simulator Demonstration Model and the 7.5% F-15 Inlet/Airframe 

Model (References 6 and 7) The tare force of this seal depends 

upon the pressure differential across it and the relative deflection between the metric and non-metric duct on either side of it 

(imposed by model aerodynamic loads) . Successful calibrations of 

these effects fran the two referenced programs provide experience 

for establishing this tare force. 

Two design features have been incorporated to further ensure 

minimum, repeatable corrections due to the duct seal tare. One 

is to minimize deflection at the duct metric break, and the other 

is to ensure identical seal positioning in all test modes. The 

latter feature is provided by the mounting arrangement, as 

addressed in Section 4.1. To minimize the vertical downward offset, which results fran balance deflections at high a (large pitching moments), a .OS0 inch upward vertical offset is incorporated 

in the basic design. The variation in offset with angle of 

attack will thus be as shown in Figure 4-3. At low a, where the 

air loads produce very little deflection, a small upward offset 

(.OS in. maximum) will exist. As the model is pitched to nominal 

maneuvering a (10°-12') and additional deflections occur, near 

zero offset will occur. At even higher a test conditions, a 

small negative downward offset (.OS in maximum) will exist 

between the inlet duct and simulator. The maximum offset over 

38 
METRIC INLET DUCT r rSlMULATOR 

(NON-METRIC) 

W-m-Ejlt 

0.050 

t --e IN. --+ 0.050 

f 

IN. --# 

LFLEXIBLE 

SEAL 

LOW a 

+0.0!50 IN. OFFSET 

INCORPORATED 

(ZERO INLET DUCT 

DE FLECTl ON 1 

- HIGH a 

(0.100 IN. INLET DUCT 

MEDIUM a 

(0.060 IN. INLET DUCT 

ZERO OFFSET -0.050 IN. OFFSET 

DEFLECTION) DEFLECTION) 

Op13.MOS.P 

FIGURE 4.3 

DEFLECTION VARIATION AT DUCT METRIC BREAK 

39 
the test angle-of-attack range is therefore limited to the .050 

inch deflection. A deflection of this same value, experienced on 

previous models, caused no problems. 

4.3 MODEL SUPPORT/TUNNEL INSTALLATION - The emphasis of this program is to evaluate the airframe/propulsion system interactions 

in the conventional wing &rne flight mode. An installation in 

the Ames 11-foot Unitary Plan Wind Tunnel has been defined to 

conduct this evaluation. There is also interest in assessing low 

speed interaction effects at angles of attack up to 90*, where 

vectored thrust must be used for control. Testing at these conditions would be conducted in a future test program in the Ames 

12-Foot Pressure Wind Tunnel. A conceptual installation for the 

12-foot tunnel testing has also been defined. 

4.3.1 .Utes 11-Foot Tunnel Installation - The support system 

was designed consistent with (1) a test Mach number range from 

0.3 to 1.4, (2) a constant tunnel Reynolds number of 2.5 x 106 

per foot, (3) angle of attack capability from -5' to +25', and 

(4) minimum buoyancy effects from the support system. 

The test Mach number range covers the operating range available in the 11-foot tunnel. The test Reynolds number was 

selected as a compromise between the desire to operate at the 

highest dynamic pressure possible and model and support system 

strength considerations. The tunnel dynamic pressure and total 

pressure corresponding to this Reynolds number is shown in Figure 

4-4. Angle of attack capability up to 25' should cover the 

conditions of primary interest for conventional V/STOL flight 

attitudes. 

- 

A tunnel installation was defined consistent with these 

operating requirements. The installation, Figure 4-5, consists 

of the Ames 40 inch extension and 10' canted adapter, and a new 

tapered adapter, offset adapter, and sting support. The maximum 

bending moment for the installation is within the 800,000 in-lb 

moment limit of the tunnel sector. 

40 
TUNNEL 

TOTAL 

PRESSURE, 

pto 

PSlA 

TEST 

SECTION 

DYNAMIC 

PRESSURE, q 

PSF 

Conditions set for 

Max P, of CMAPS-Re = 2.5 x lOC/ft 

FIGURE 44 

PLANNED TEST OPERATING CONDITIONS IN NASAlAMES 11 FT UNITARY TUNNEL 

41 
42 
The model sting length was selected to minimize the buoyancy 

i effect of the offset adapter on the model. The effect of the 

I adapter was analyzed using a MCAIR-developed program for analyz- l 

ing subsonic and transonic flow about 3-9 bodies (Program 

AFTEND). The analysis was performed for two different sting 

lengths. The shorter sting length is compatible with both the 

ll-foot tunnel installation and high angle-of-attack testing (40' 

to 90") in the 12-foot low speed tunnel. The longer sting represents an estimate of the required length to minimize the effect 

of the adapter, as determined fran AEDC sting interference test 

results, Reference 8 For comparison, a baseline sting installation without the offset adapter was also analyzed, Results are 

shown in Figure 4-6 in terms of incremental pressure coefficient 

(bCp) from the baseline for the short and long stingsThe offset adapter with the short sting produces a Cp rise 

of zp=rsxL~ztely 3.02 on tne model aftbody up to the metric 

break, which converts to a drag error for both nozzles of about 6 

counts. In this calculation, the -02 ACp value was applied to 

the projected area fran the aft metric break to the nozzle exit. 

The Cp increase with the short sting was considered unacceptable. 

There is no discernible pressure rise from the adapter felt on 

the model with the longer sting. Based on this analysis, and in 

concurrence with the NASA project engineer, the longer sting has 

been selected for the program. 

4.3.2 Ames 12-Foot Tunnel Installation - NASA-Ames is currently planning to modify the 12-foot tunnel: the extent of this 

modification is not totally defined. However, a conceptual 

installation has been defined using the current AMES pitch system, Figure 4-7. With this installation, angle of attack testing 

up to 90' would be conducted in two steps. For angles of attack 

from 0" to 40°, the installation consists of an existing NASA 

adapter, a new tapered adapter, and the 0" offset adapter used in 

the ll-foot installation. For the testing to 90' angle of 

attack, another tapered adapter will be provided between the 0' 

43 
MODEL 

BASELINE (NO ADAPTER) 

/ 

SHORT STING 

LONG STING 

MODEL STATION on-u - IN. 

FIGURE 4-6 

STING LENGTH EFFECT ON ADAPTER INDUCED PRESSURE INCREMENTS 

44 
i 

0 

.’ 

s 

i 

T 

% 

I 

Y 

3 

a 

0 

0 

a 

LL 

a 

Y 

z 

z 

3 

c 

W 

pc’ 

v) 

a 

a 

8 

P 

c 

IL 

c! 

45 
offset adapter and the Ames high angle of attack pitch system. 

It should be noted that the model sting shown for the 90' 

testing is shorter than the sting seiected for the 11-foot 

instailation. 

4.4 ,FORCE 3AwNCZ - A Task Mar!< XXXII, 2.5 inch balance owned by 

NASA will be used in this program. Its selection was based on 

maximizing the axial force accuracy, consistent with normal 

force/pitching moment and deflection limits. The axial force 

limit of the balance is 250 lb, which is consistent with the 

expected maximum axial force loads for the testing (near 200 lb 

at Mach 1.4, maximum a (16')). The maximum allowable normal force/ 

pitching moment envelope for the Mark XXXII is compared to the 

estimated model loads in Figure 4-8 for Mach 0.9 and 1.4. Also 

shown are the allowable deflection limits at the duct and aft 

metric breaks, AY = 0.100 and 0.080 inches respectively. The 

deflection limits are based on test experience from applicable 

inlet and nozzle research programs, References 5 and 6. At Mach 

0.9, the baseline model configuration testing at the higher 

angles of attack ( a>2O0) - will be limited to negative canard 

deflection settings (canard leading edge down), due primarily to 

the duct metric break deflection limit. At Mach 1.4, the balance 

load capacity limits the test angle of attack as a function of 

canard deflection. 

The available test envelope is large enough to cover the 

conditions of primary interest. The angle of attack/canard 

deflection test envelope is summarized in Figure 4-9. 

4.5 SALANCE THERMAL CONTROL - The force balance in the model 

will be subjected to a wide range of temperatures in the three 

test modes. In the simulator mode, the WS units have three 

significant heat sources: (1) turbine drive air with a fixed 

temperature of about 200°F, (2) turbine bleed air with a 

temperature varying from about 50' to 165'F, and (3) compressor 

46 
Mo = 0.9: q = 450 psf; Bat G = MS 40 - 

I3 

4' 

0 

4 

-8. 

BALANCE 

NORMAL 

FORCE 

1,Ooo LB 

I I AY = 0.700 IN. 8 DUCT 

METRIC 8REAKAY = 0.080 IN. @ AFT 

-2.5 IN. MK XXXlI 

BALANCE ENVELOPE 

M0=1.4; q=550 psf; Bal Q =MS 40 

8- 

4 

BALANCE 

NORMAL 

FORCE 0 

1.000 LB 

' 

4 

-8 - 

-40 -30 -20 -10 0 10 20 30 

BALANCE PITCH MOMENT - 1,000 IN.-LB 

QPlSOEO3-17 

Note: (1 1 Canard "Off:' Simulates Ultra CloseCoupled Configuration 

FIGURE 44 

MODEL LOADS AND BALANCE NORMAL FORCEIPITCHING MOMENT ENVELOPE 

47 
0 BASELINE CONFIGURATION 

M, - d9;q = 45OP8f 

20 

10 

CANARD 0 

DE FLECTl ON 

ANGLE, 6, 

DEG 

-1 0 

-20 

-30 

0 10 20 30 40 

ANGLEOF-ATTACK, u - DEG 

M, - 1.4 q = 550 psf 

r CANARD DEFLECTION LIMIT 

I I 

0 ULTRA CLOSECOUPLED CONFIGURATION (CANARD "OFF") 

0 10 20 30 40 

ANGLEOF-ATTACK, a - DEG 

FIGURE 4-9 

DEFLECTION TEST ENVELOPE 

2.5 in. MKXXXll Balance Bal = FS 40 

ANG LE-0 F-ATTAC KICAN AR D 

48 
discharge air temperature that can approach 400'F. In the jeteffects mode, turbine drive and bleed air temperatures of the 

simulator mode are duplicated, although the high temperature 

compressor discharge air is not. In the flow-through mode, the 

balance is subjected only to the tunnel temperature ('L 120'F). 

Bias errors can be introduced into the force balance output as a 

result of the differing temperature environments, particularly if 

thermal gradients occur. 

These errors can be eliminated if the balance is thermally 

controlled in all modes to the same constant temperature levels, 

with minimum thermal gradients. A three-dimensional thermal 

model of the balance and surrounding model structure was developed as a means to design the thermal control system. The model 

consists of approximately 300 nodes: a schematic of the node 

locations are shown in Figure 4-10. 

The thermal control system was designed under the following criteria: 

(1) Maintain a balance operating temperature of 160'F in 

all test modes, consistent with the nominal upper limit 

of Task balance operation: 

(2 Minimize steady-state balance axial temperature gradients to within 2'F, consistent with recent experience 

on an F-15 jet-effects model: and on the ANC program; 

(3) Provide steady state temperatures in a reasonable 

amount of time (% 1-2 hr). 

The thermal control system designed to these criteria is shown in 

Figure 4-11. Two heaters (#1 and #2) are located on the forward 

and aft sections of the balance housing. Another heater (#3) is 

located on the forward end of the balance taper insert. The 

fourth heater (#4) is located beneath the balance. Very thin 

49 
d 

n 

- 

I50 

2 

-0 

E 

B 

u) 
51 

w 
( .030 inch) custom heaters, manufactured by the Watlow Corporation, will be used for this application. Two thermocouples are 

provided for each of the four heaters. One provides feedback to 

the proportional temperature controller, set to naintain 160 "F on 

a specific location of the balance housing (or tapered insert). 

The other is for temperature monitoring and back-up purposes. 

Use of the proportional temperature controllers, in conjunction 

with the thin blankets, will permit heater power densities of up 

to 25-30 watts/in2 for this application. 

The thermal model was used to predict balance temperatures 

with this thermal control system for both transient and steadystate operating conditions. The transient analysis, representing 

balance heat-up from 80.F with the tunnel off, indicated the 

balance could be driven to 160'F in about one hour. This is consistent with the design criteria. Maximum power density required 

for this length of heat-up time was about 1/3 of the system's 

capability. This converts into either a high safety factor or a 

potentially more rapid heat-up period. The power density requirements for steady-state operation with the model and tunnel operating are much less. 

Three steady-state cases, two for the simulator mode and one 

for the flow-through mode, were analyzed to ensure that the ternperature gradient requirements were attained. The two simulator 

test cases correspond to the range of CMAPS turbine bleed air 

temperature (50 to 165.F) from recent CMAPS testing. A maximum 

axial temperature gradient of about 2.F is predicted for the 

three test cases, as shown in Figures 4-12 and 4-13. This 'is 

within acceptable limits. The angular and radial temperature 

distributions for these same three conditions are within about 

4'F, as shown in Figure 4-14. 

Detailed 

trollers will 

activity. 

design and procurement of the heaters and conbe accomplished in the Phase 2 detailed design 

52 
LOWER BALANCE ‘-INNER BALANCE 

SLEEVE ROD 

BALANCE 

STEADY 

STATE 

TEMPERATURES 

OF 

I I I I 

INCHES AFT FROM NOSE OF BALANCE 

4 8 12 16 20 

OP1~ 

FIGURE 412a 

BALANCE AXIAL TEMPERATURES - SIMULATOR MODE TBLEED = 50°F 

53 
I 

I 

‘L INNER BALANCE 

ROD OUTER BALANCE 

SLEEVE 

0-0 Outer balance sleeve 

loo 

160 

4 8 12 16 20 0 

150 I 

140 

BALANCE 

STEADY 

STATE 

TEMPERATURES 

130 OF 

1 20 

110 

I 

SIMUl.!TOR MODE 

MACH 1.4 

T - 2WoF 

P= 275 PSlA 

m = 1.0 LBM/SEC 

- DRIVE AIR 

BLEED AIR 

- T-leF 

P-220PSlA 

m = 0.5 LBMJSEC 

54 
I 

170 

160 

150 

140 

BALANCE 

STEADY 

STATE 

TEMPERATURES 

OF 130 

1 20 ? 

110 

b 

r 

4 8 12 16 20 

loo 0 

55 
OUTER BALANCE 

SLEEVE 

SIMUUTOR MODE 

I yl60.10 

160.66 

160.94 

Ll61.62 ' L63.48 

161.16 

160.26 

160.96 

Ll64.84 

SECTION A-A SECTION 6-6 SECTION C-C 

.- 

I f 159.80 

160.18 

160.40 

' .\-160.86 

SECTION A-A 

FLOW 

__-_ 

I r159.S. 

159.63 

159.69 ' 160.52 

SECTION A-A 

I f 159.87 

159:#1: 159.98 

' L160.84 

158.65 

5838 158.54 g158.74 ' 157.00 

SECTlON 6-8 

P 

I r160.0

159.32 159.33 #I 

I L160.37 

._ SECTION C-C 

12OOF) . 

I r159.54 

158.79 

158*63#~5~.~ 158.1 1 

L155.75 

SECTION C-C 

FIGURE 4-14 

SALANCE ANGULAR AND RADIAL TEMPERATURES GPlSOSOS7 

56 
I 

4.6 INLET DESIGNS - Two different inlets have been designed for 

the baseline and ultra close-coupled configurations. 

4.6.1 Baseline Inlet - The baseline inlet is a rectangular, 

normal shock design with a 15' scarf angle. The inlet concept is 

shown in c"igure 4-15. It was sized for operation at 95 percent 

of critical mass flow ratio with the maximum simulator airflow 

(1.65 lbs/sec). A rotating cowl lip is incorporated into the 

design to reduce lip separation at static conditions and at high 

angles of attack. Lip rotation angles of 45' and 70' were 

selected for this design. It is expected that the 45' configuration will be used for low speed at moderate to high angles of 

attack. The 70' cowl deflection would be used for static operation and at low speed, high angles of attack (approaching 90'). 

Sideplates will be incorporated when the cowl is deflected. In 

addition, the cowl rotation point has been selected to provide a 

Parge flap knee radius to prevent separation. The rotating cowl 

lip design for this inlet is based on a similar design which was 

recently tested at MCAIR. 

The need for specialized inlet devices, such as the rotating 

cowl lip, is illustrated by the test results on the 8.5% Simulator Demonstration Model, Reference 6. In this testing, there 

were numerous instances of simulator compressor stalls at static 

operation which were attributed to high flow inlet distortion 

levels. This distortion was due to separation off of the relatively sharp (F-15 type) lip at static conditions. The problems 

were resolved by subsequently using a bellmouth adapter as shown 

in Figure 4-16. 

For this program, the rotating cowl lip was selected because 

of its demonstrated performance in the Reference 9 investigation 

and also in a recent MCAIR Independent Research and Development 

(1-1 test program. me rotating cowl lip configuration tested 

in the recent McAIR program is shown in Figure 4-17. The inlet 

performance improvement with the rotating cowl lip configurations 

57 
A I 

ROTATING COWL 

SIDEPLATE A 

COWL LIP ROTATION 

POINT 

i 

ENGINE FACE 

WL 10.64 

MODEL LWR NACELLE ML 

QPlSOSOWO 

FIGURE 4-15 

ARTICULATING COWL LIP FLAP SYSTEM FOR BASELINE INLET CONFIGURATION 

58 
r 

Inlet with Adapter 

To P 

-- 

Looking Aft 

Bm Inlet 

Lines of Constant 

Looking Aft (2% From Average) 

om1~ 

Figure 4-16. 

INLn STATIC PERFORMANCE IMPROVEMENT WITH BELLMOUTH FOR 

8.5% SIMULATOR DEMONSTRATiON MODEL 

59 
FIGURE 4-17 

MCAIR ROTATING COWL LIP CONFIGURATION 

60 
was determined statically and at Mach 0.25 over an angle of 

attack range Of 0' to 90'. As shown in Figure 4-18, the rotating 

cowl lips provide a significant performance improvement. 

A caparison was made of inlet recovery, steady state distortion, and turbulence at static conditions for the following 

configurations: 

o 8.5% Sim. Dev. Baseline Inlet (F-15) 

o 8.5% Sh. Dev. Bellmouth Adapter 

0 MCAIR 40' and 70' Rotating Lip Configurations 

The performance differences are tabulated in Figure 4-19. Note 

that the MCAIR 70. rotating lip inlet has lower distortion than 

the bellmouth inlet and only slightly lower pressure recovery. 

Shcc tkt V/STGL isEating COW1 lip design are geometrically 

similar to those recently tested at MCAIR, the V/STOL inlet 

should have comparable performance and therefore result in 

satisfactory CMAPS operation. 

An evaluation was also conducted to determine the need for 

an inlet bypass system to maintain stable inlet operation when 

the simulator is windmilling at supersonic test conditions. The 

operating characteristics at Mach 1.4 of a noma1 shock inlet 

similar to that on the baseline V/STOL configuration are shown in 

Figure 4-20. The start of inlet flow instability or '*buzz'* 

occurs below the mass flow ratio associated with the simulator 

windmilling airflow for all positive angles-of-attack tested. A 

bypass system is therefore not necessary providing the model is 

positioned at positive angle of attack during windmilling 

operation. 

4.6.2 Ultra Close-Coupled Inlet Definition - The conversion 

of the baseline inlet configuration to a modified ultra-closecoupled configuration is accomplished with a minimum of hardware 

61 
leer 

PRESSURE 

RECOVERY, PaIPa 

0.8 L 

a4r 

STEAOY STATE DISTORTION ["* rn;;;:) mi"] 0.1 1 

TURBULENCE OT 

0 L 

STAT1 C 

OPERATION 

M, - 0.25 

J = 00 

Notes: 

-1 

B 

0 

B 

0 

0 

0 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I \ 

70° LIP 

400 LIP 

BASELl N E 

BASELINE 

400 LIP 

70° LIP 

BASELl N E / 

J /400 LIP 700 LIP 

I I I I I 1 

0 20 40 60 80 100 

ANGLE-OF-ATTACK, OL 

QPl3.OW3.S7 

1. Data from MCAlR IRA0 test program. 

2. Apt rms is the root mean square deviation of the fluctuating portion of the 

dynamic total pressure. 

FIGURE 4-18 

IMPROVED INLET PERFORMANCE WITH ROTATING COWL LIP 

DATA AT DESIGN ENGINE AIRFLOW 

62 
Performance Parameter 

Inlet Recovery 

Steady State Distortion 

Turbulence 

Figure 4-19. Inlet Performance Comparison 

Static Operation 

lnht Configuration 

8.5% Sim bv 8.5% Sim Dev MCAlR 

(See Note 21 , ~~apter 

Betlmouth 70' Rotating F-15 

Inlet Lip 

0.870 0.881 0.960 0.933 

0.138 0.1 27 0.104 0.071 

N/A 0.034 N/A 0.01 5 

2 

z 

0 0.7 

Y) 

- - m 

e 

1 oo 

- 

0.6 I I I I I I I I I 

0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 .o 

Mass Flow Ratio, Ao/A, 

GP1uyOl4 

Figure 4-20. Typical Normal Shock Inlet Stable Range 

63 
changes. As illustrated in Figure 4-21, this modification 

involves ' (1) remoGal of the canard, (2) removal of the basic 

inlet duct, (3) insertion of a modified inlet duct, and (4) modification of the inner wing. All other aspects of the model 

hardware remain undisturbed as this modification is made. The 

balance and seal tare calibrations thus remain the same. 

The short nacelle necessitates changes to the basic inlet 

design concept. In keeping with the optional design guidelines 

established by NASA, a subsonic inlet design has been selected, 

as detailed in Figure 4-22. Due to the short duct length, a circular inlet shape was chosen to avoid relatively large diffuser 

wall angles which would result when transitioning from a square 

to a circular duct over a very short distance. Based on similar 

designs, the inlet was sized to an inlet throat Mach number of 

0.55 at the simulator design airflow. 

The articulating lower cowl flap used in the baseline inlet 

design for improved performance at static and high angle of 

attack conditions is not being used in the short nacelle inlet. 

It was felt that the articulating flap was not compatible with 

the circular subsonic inlet concept. Therefore, three other 

design features are being used to "replace" the articulating flap 

function (good static/high angle of attack Performance) . First, 

in accordance with the results from Reference 10, a reverse scarf 

(lower lip extending past the upper lip) is incorporated. Tne 

. increase in angle of attack capability obtained with a reverse 

scarf inlet design is shown in Figure 4-23. Second, generous 

inlet contraction ratios (capture area divided by throat area) 

have been selected for this inlet design. Based on the YC-15 

design, contraction ratios of 1.78 and 1.34 have been chosen for 

the lower and upper inlet lips, respectively. The angle of 

attack performance improvement with increased inlet contraction 

ratio is shown in Figure 4-24. Finally, a leading edge slat has 

been provided over the lower 120" of the inlet to help prevent 

separation on the lower lip. AS presently ,lanned, the slat will 

64 
1 /I 

ULTRA 

CLOSE-COUPLED / 

iNLET- I I1 A ,/I __ A~BEN NOZZLE-’ 

WL 10.640 -1 

BL 6324 / INNER WING FAIRING^ (FS 32.2 

. I 

WING -- NOT SHOWN 

Note: This arrangement is applicable LLEADING EDGE SLAT 

to all test modes. 

FIGURE 4.21 

MODEL ARRANGEMENT OF CLOSEGOUPLED CONFIGURATION 

65 
ML PER YC-15 DESIGN 

MATE TO 

EXIST1 NG M. 

t 

3 IN. 

CMAPS DIA 

1 

BL 6.324 

Q WL 10.64 

A 

MATE TO Ii EXISTING ML 

+ +CONSTANT 

t AREA SECTION 

I 

6 FS 34.50 2:1 FS 30.65 FS 31.90 FS 34.50 

LEADING EDGE 

LIP SLAT 

63 

DETAIL "B' 

View A-A 

- - -- BL 6.324 

21 ELLIPSE 

TYP 2 PLACES 

FS 34.50 

FIGURE 4-22 

ULTRA CLOSE-COUPLED INLET CONFIGURATION 

66 
A - Reverse Scarf 

a 

g: 

P 

U ii 80- 

0 - Baseline 

Mo = 0.12 

Design M, 

30 

0.3 0.4 0.5 0.6 0.7 0.8 

Throat Mach Number, Mt 

a GP130401.7 

Figure 4.23. Effect of Reverse Scarf lniet on Flow 

Separation Bounds 

67 
" 

1.0 1.2 1.4 1.6 1.8 2.0 2.2 

CR - Inlet Lip Contraction Ratio 

QPl3-04014 

Figure 4-24. Effect of Inlet Contraction Ratio 

on Flow Separation Bounds 

68 
be used only at very high angle of attack. Performance improvement with a geometrically similar slot design is illustrated in 

Figure 4-25. 

It is felt that this ultra close-coupled inlet design will 

provide the necessary performance levels at static and high angle 

of attack conditions to ensure safe simulator operation. 

4.7 NOZZLE DESIGNS - The basic nozzle design for the program is 

the GE ALBEN concept. The ALBEN meets the program requirements 

and was selected for cost considerations since it is existing 

hardware fran the ANC program. It can be tested in any of the 

three modes, although it will be used primarily in simulator and 

jet-effects testing. Nozzle extensions are used for the reference nozzle condition and in the flow-through mode. The aerodynamic design of the nozzle extensions is discussed Lc Sectis: 

3: their use in the flow-through mode is discussed in Section 5. 

The ALBEN nozzle assembly simulates a full scale GE design 

with internal/external expansion, an aspect ratio of 3.8, fully 

variable throat area control, and thrust vectoring capability. 

This is a specialized type of single expansion ramp nozzle 

designed particularly for low weight. It utilizes an elliptically shaped canister or shroud to vary throat area, with the 

bottm section forming the lower nozzle lip. The upper portion 

of the shroud surrounds the upper ramp assembly creating a cavity 

that is pressurized with fan air. The pressure loading on the 

upper section of the shroud essentially balances the load on the 

lower section. 

The ALBEN model can be vectored both at dry power and afterburning to 20' and 30° deflectim. T3e x&ei representation of 

the ALBEN is shown in Figure 4-26. 

69 
CR = 1 .J6; With Slai 

CR = 1.46; No Slat 

I 1 I 1 

0 2 4 6 8 10 01 

Throat-to-Freestream Velocity Ratio, Vt/Vo op,to40,~, 

Figure 4.25. Effect of Leading-Edge Slat on Flow Separation Bounds 

FS 50.689 

I 

-- 

DRY POWER 

----- = 6.481 IN.2 

ATh = 5.151 IN.' AeXaxt = 7.307 IN.2 

LOW MACH AI0 

OC13.06SSM 

. FIGURE426 

GE ALBEN NOZZLE ASSEMBLY 

AR = 3.8:l 

70 
5.0 DESCRIPTION OF TEST MODE CONCEPTS 

Turbine inlet flow (W4) = 5.5 lbs/sec/CMAPS unit 

Turbine inlet pressure (PT4) = 1300 psia 

Turbine inlet temperature (TT4) = 200°F 

Compressor inlet pressure (PT2) = 16 psia 

Successful completion Of program objectives depends upon 

attention to detail in the design of all three test modes. 

Zonceptual layouts and associated analyses have Seen accomplished 

for the key design aspects of the simulator, flow-through, and 

jet-effects testing modes. 

5.1 SIMJLATOR TEST MODE CONCEPT - The simulator mode is the most 

complicated of the designs, and requires extensive analysis to 

ensure the model and simulator are compatible. The only hardware 

components unique to this mode are the actual CMAPS units. The 

layout of the simulator test mode concept is shown in Figure 5-1. 

Extensive analyses of the pressure loss characteristics of the 

drive and bleed air system were conducted based on this design, 

and a "flexibility" analysls was &r,e to ~ssess CiSjmodei compatibility. The term flexibility refers to the CMAPS capability, 

in the model, to provide independent variation of airflow or 

engine pressure ratio while holding the other parameter fixed. 

5.1.1 Drive Air System - The turbine drive air system.must 

provide the simulator the required flow at a given pressure 

level. The design conditions for the drive air system are at 

maximum flaw and pressure, as determined from data of References 

(6) and (11). 

The proposed drive air system is composed of three different 

sections. A schematic is shown in Figure 5-2. The system consists of a 1-inch I.D. line from the NASA Flow Control Pallet to 

the model strut. Once inside the strut, the line size is reduced 

71 
72 

i 
rl 

r-- '-. .. - 

i 

I 

I_ .i 

73 
to a .683 inch I.D. line (flow area of .366 id) to minimize 

strut size. The drive air line area is increased to .70 in2 

inside the model as soon as possible to reduce pressure loss. 

This .70 in2 area is maintained up to the uWS drive manifold. 

It should be noted that this larger flow area is about the same 

.as the area at the flange of the current WS drive manifold 

shell. 

Assuming an available supply pressure of 2000 psia at the 

exit of the NASA Flow Control Pallet, sufficient pressure should 

be available to the CMAPS units. With 2000 psia supply pressure, 

the available pressure at the CMAPS turbine inlet is estimated to 

be 1575 psia (1300 psia maximum required). The complete loss 

analysis is contained in Appendix A. 

5.1.2 Bleed Air System - The capability of the CMAPS to 

vary engine pressure ratio (EPR) at a fixed compressor airflow is 

provided by mixing in the nozzle all or part of the turbine bleed 

airflow with the compressor airflow. For a given nozzle throat 

area and mixer area, the pressure drop or flow restriction in the 

bleed system directly affects the range of EPR available. As 

bleed system pressure losses increase, the amount of bleed flow 

that can be handled before choking occurs is correspondingly 

reduced. Consequently, the rest of the bleed flow must be dumped 

through the mixer, which in effect raises the minimum EPR attainable at a given compressor airflow. Larger reductions in bleed 

line area (or increased pressure loss) eventually result in a 

condition where it is no longer possible to power the CMAPS at 

maximum speed without stalling the compressor. Therefore, the 

design of the turSine bleed air line system is critical since it 

has a direct impact on the CMAPS performance. This is especially 

true for small throat area nozzles (typical of engine dry power 

operation), where the majority of the turbine bleed air must be 

routed back through the bleed air line. 

74 
To provide the greatest CMAPS flexibility, the bleed system 

was designed with the largest line areas possible considering the 

model and support system constraints. In the design shown in 

Figure 5-3, a bleed Line area of 1.05 in2 is inaintained inside 

the model between the CMAPS bleed manifold and support strut. 

Once inside the forward section of the strut, the bleed area is 

increased to 1.23 in2 to decrease pressure loss. The bleed area 

is further increased to 1.62 in2 in the aft section of the strut. 

Aft of the model strut, a 2.0 in I.D. flexible line ducts the air 

to the NASA Flow Control Pallet located on the tunnel traversing 

strut. The bleed air then passes through the flow control/ 

measuring hardware on the pallet, and is exhausted into the 

tunnel. 

With the system thus defined, the pressure drop through each 

component was calculated. A loss sumnary is shown in sumnary 

form in Figure 5-4. The bleed system losses are separated into 

three categories: (1) losses between the bleed manifold and 

model strut, (2) losses in the model strut, and (3) losses 

through the Flow Control Pallet hardware. About 60% of the total 

pressure loss occurs inside the model, where line size is 

restricted due to space limitations. An additional 20% is lost 

in the model strut, -where strut size is a controlling consideration. Finally, about 20% of the pressure is lost within the 

pallet installation. The pressure loss analysis is described in 

Appendix B. 

5.1.3 Flexibility Analysis - Extensive study has been 

conducted to estimate the flexibility envelope of the CMAPS units 

in the model. This work was done in cooperation with General 

Electric. The key "independent variables" which affect flexibility are: 

(a) Turbine bleed system pressure loss 

(b) Bleed venturi area (Av) 

(c) CMAPS mixer area (~~57) 

(dl Nozzle throat area (Ag) 

75 
7- 

I 

i 

I 

1 I 

I 

i 

76 

-- - \ 

\ 

\, 
I 

/I 

_- .---- -- -_ > 

2 

I 

NO. 1 

REGION NO. 2 REGION NO. 3 

i 

REGION % OF TOTAL A P, LOSS 

NO. 1 INSIDE MODEL.. ......... 62% 

NO. 2 MODEL STRUT.. ......... 18% 

NO. 3 PALLET INSTALLED.. .... 20% 

QP1~58 

FIGURE 54 

TURBINE BLEED SYSTEM PRESSURE LOSS SUMMARY 

77 
In the detailed initial analysis presented in the March 1981 

progress report, each of the above parameters was varied independently to assess their affects on CMAPS flexibility. Results 

indicated that reduction of bleed system 9ressure Loss should 5e 

further evaluated, as well as a more optimum (smaller) mixer 

area.. It was also determined that the larger of the two available venturis should be used. An increase in nozzle throat area 

was found to be effective, but was not considered due to the cost 

of new hardware. 

Subsequent analyses has led to the final recommended system 

to achieve good CMAPS/model compatibility in terms of the flexibility envelope. The basis for comparison is the CMAPS flexibility obtained during the recent simulator development test at 

AEDC, Reference 1. A comparison of the key flexibility 

parameters for the CMAPS development system and the recommended 

NASA system, at the critical dry power condition, is given below: 

Parameter 

Turbine Bleed System Total 

Pressure Loss 

CWS 

Development Recommended 

system NASA System 

- 20% (Bleed 42% (Bleed 

Manifold Manif old 

to Venturi) to Venturi)* 

Bleed Venturi Area, A, (in2) 0.65 0.85 

Dry Power Nozzle Throat Area, 3.098 2.83 

Ag (in2) 

CWS Mixer Area, AE57 (in2 ) .lo73 0.062 

*Bleed air system defined in paragraph 5.1.2 

78 
The recommended parameter values above represent ( 1) an 

increased model bleed line area, compared to the earlier design, 

of about lo%, (2) the use of the larger of the two available NASA 

venturis, (3) reduced area mixer and (4) the use of the currently 

available ALBEN nozzle from the ANC program. The minimum CMAPS 

operating line for the recommended NASA system is shown in Figure 

5-5, compared to the CMAPS development system. This prediction 

has been made using the GE CMAPS cycle analysis and the noted 

inputs. The recommended system is considered well within acceptable limits, with capability to achieve maximum CMAPS airflow 

(1.65 lb/sec) with adequate stall margin. 

5.2 FLOW-THROUGH TEST MIDE CONCEPT - The testing in the conventional flow-through mode is designed to measure the basic vehicle 

aerodynamics, including canard trim and inlet drag characteristics and any associated interactions. Conceptual design of the 

flow-through mode is shown in Figure 5-6. 

The only model change required to convert from the simulator 

to the flow-through mode is to remove the CMAPS units and install 

the flowthrough duct. The duct includes a dummy hub to simulate 

the simulator compressor hub. The nozzle extensions are normally 

installed in the flow-through de, although the ALBEN can be 

tested if desired. The flow-through duct is mounted using the 

same attachment as the CMAPS to ensure identical positioning of 

the inlet duct seal. The nozzle extensions are cantilevered from 

the common transition duct/nozzle support. 

Inlet mass flow ratio is controlled with chokes at the 

nozzle extension exit. A set of four chokes is sufficient to 

cover the range of inlet mass flow ratios available with the 

CMAPS units. The choke exit areas will range from approximately 

2.5 to 5.2 in2. Chokes of similar areas were utilized in the 

8.5% Simulator Development Program and effectively covered the 

airflow range of interest. At Mach 0.6 and below, an ejector 

will be required to obtain the full range of inlet mass flow 

79 
ENGINE 

PRESSURE 

RATIO 

PT8/PT2 

COMPRESSOR CORRECTED AIRFLOW y)t2 - 

FIGURE 5-5 

ESTIMATED SIMULATOR FLEXIBILITY ENVELOPE 

M, = 0.3; PT2 = 15.2; P,,b = 15.1 

80 
81 
ratios. It will be mounted on the tunnel support, as shown in 

Figure 5-7. . Flexible lines will connect the ejector to the 

model. The CMAPS air supply will be adequate to power the 

P j ector . 

The recommended ejector for the lower speed testing is XAIR 

owned and was previously used at NASA/Ames in an inlet development program, Reference (12). In this program, the ejector was 

used to pump up to 7.0 lbm/sec corrected airflow. The ejector 

flow/pressure requirements are presented in Figure 5-8. 

5.3 JET-EFFECTS TEST MODE CONCEPT - Testing in the conventional 

jet-effects mode is designed to measure the incremental aerodynamics effects due to nozzle geometry (dry and afterburning), 

nozzle pressure ratio, and vectoring. Conceptual design of the 

jet-effects mode is shown in Figure 5-9. 

The only model change required to convert frm the simulator 

to the jet-effects mode is to remove the CMAPS units and install 

the high pressure air supply plenum assembly. Inlet fairings are 

also added. 

In this mode, the non-metric plenum assembly is attached 

using the identical attachments as for the other two modes. The 

duct metric break seal and instrumentation ring forward of the 

plenum are retained, even though they do not serve a function. 

This is to ensure consistency in the duct metric break seal tare 

among the test modes. The bellows and transition duct aft of the 

plenum are also retained. The &BEN nozzles mount to the transition duct/nozzle support, thus vectoring loads are transmitted 

directly to the balance housing. 

The high pressure jet-effects plenum (Figure 5-9) is 

designed to provide the same total pressure levels and distribution at the nozzle charging station as the CMAPS. To accomplish 

this, the actual WS mixers will be utilized as part of zhe air 

82 
CMAPS AIR CONTROL SYSTEM 

1.25 DIA FLEXIBLE 

REFERENCE NOZZL 

EXTENSIONS (2 PLACES) 

CMAPS AIR CONTROL SYSTEM 

1.25 DIA FLEXIB 

TUNNEL 

FLEXIBLE AIR LINES; 

CHOKE-TO-EJECTOR 

Q 

.- 

-I 

OF ROTATION 

TS 304.75 

I if 

-- I MCAIR EJECTOR J ! 

I 

I TS ST'ATIOI 

an-ar 

WINDOW Q 

FIGURE 5-7 

EJECTOR INSTALLATION FOR LOW SPEED FLOW THROUGH TESTING 

83 
EJECTOR 

PRIMARY 

FLOW TOTAL 

PRESSURE 

PSIA 

INLET CORRECTED AlRF LOW - LSM/SEC Qn-, 

FIGURE 5-8 

EJ ECTO R FLOW / PRESSU RE R EQ U I REM ENTS STATIC OP ERATlO N 

Static Condition 

84 

EJECTOR 

PRIMARY 

MASS FLOW 

LBIVI/SEC 
85 
supply plenum. Two separate high pressure air supplies are provided to the plenum. One supply Comes in through the same line 

that provides the drive air to the CMAPS, and thus represents the 

compressor airflow. This flow passes through a choke plate and 

in between the mixer lobes. The other supply, simulating the 

CMAPS turbine exit flow, comes through the CMAPS bleed air lines 

directly into the mixer. After passing through the mixer chokes, 

this air combines with the "compressor" air. By independent 

control of the pressure of each supply, it should be possible to 

match both the nozzle flow pressure level and distribution profile of the actual CMAPS at a variety of conditions. 

This plenum design should also match the actual CMAPS nozzle 

temperature distribution, since the actual CMAPS mixers are used. 

The current CMAPS design has demonstrated less than 6% nozzle 

temperature distortion, Figure 5-10. These data were measured at 

dry power operation during the simulator development program. 

An analysis was conducted to assess the feasibility of also 

matching the actual temperature levels of the CMAPS compressor 

exit and turbine exit flows. To match the CMAPS compressor exit 

properties, air heated as high as 400'F must be supplied through 

the model. Under this condition, the temperature of the force 

balance would be near the 180'F limit of the balance. In addition, the thermal analysis indicated a large axial temperature 

gradient (=SO'F) would occur along the balance. Therefore, to 

provide a more uniform temperature environment for the balance 

and thereby reduce bias errors between test modes, it was decided 

to. supply the jet-effects air at the same temperature levels as 

in the simulator mode, or about 200'F. Jet-effects operation at 

slightly lower nozzle temperatures than in the CMAPS mode should 

not greatly affect the data comparison between test modes, based 

on the Reference 13 test data. These data indicate only a 1 drag 

count difference due to nozzle jet temperature variations between 

500 and 8OO'R, as shown in Figure 5-11. 

86 
COMPRESSOR 

PRESSURE 

RATIO 

0.06 - 

T 7max - T 7min 

T 7avg ) T 7 Dist = x 100 

0.6 0.8 1 .o 1.2 1.4 1.6 1.8 2.0 

WA2R OP1~10 

Based on Typical Fighter Wing Area 

- 

Ref: 1980 CMAPS testing at AEDC 

wl 

E 

3. 

a 2 z .o 

'5 0.04 

:z 

gs 0.02 0- 

CI z8 

.c 

U 

0 

FIGURE 5.10 

CMAPS NOZZLE TEMPERATURE (l7) DISTORTION 

? 

0 A Nozzle Pressure 

- Ratio - NPR 

0 

Area of Interest 

I I I I 

a 

I 

I-+ - 

Figure 5.11. Effect of Jet Temperature on Afterbody Drag 

M, = 0.9 :5" Soattail Convergent Nozzle Geometry 

87 
6.0 INSTRUMENTATION 

The model and CWS instrumentation for this program are 

extensive. Model force balance instrumentation is provided, as 

well as extensive internal flow measurements and external pressure instrumentation. CMAPS instrumentation is required to 

monitor/control operation and health. All instrumentation definition was thoroughly reviewed with NASA in Phase 1 to arrive at a 

preliminary set of requirements. In addition, the instrumentation routing in the model received considerable attention. 

6.1 MODEL INSTRUMENTATION - The majority of the model instrumentation is common to all test modes. Some is test mode peculiar, relating to control/measurement of unique propulsion system 

simulation parameters, such as airflow and nozzle pressure ratio. 

A summary of the instrumentation and function is presented in 

Figure 6-1 for all test modes. 

The model surface pressure instrumentation is used to diagnose the localized effects of flowfield coupling. Locations for 

these pressure orifices is defined in Figure 6-2. Since the 

nozzle boattail is non-metric in this installation, extensive 

boattail pressure measurement is provided for integration and 

assessment of boattail drag, as detailed in Figure 6-3. 

6.2 CMAPS INSTRUMENTATION - The NASA furnished instrumentation 

for the CMAPS is used to monitor the units health and operation 

and to provide control variables. A list of this instrumentation 

is presented in Figure 6-4. 

6.3 INSTRUMENTATION ROUTING - Routing of the large amount of the 

instrumentation in this model is critical, especially in the 

simulator mode. A full size .layout of the overall model instrumentation routing is provided as Enclosure (2) to this repkt 

transmittal. The model static pressures will be routed to 

88 
3 

c 

0 

c 

0 

e 

v) 

I 

c 

0, 

a 

2 

._ 

a 

a 

3 

w 

t; 

1 

e 

0 

0 

c 

v) 

a 

a 

I 

c 

z 

0 

a 

s 

.. 

d 

c v) 

-1 

0 

a 

z 

e 

0, 

a 

3 

a 

ry 

c 

w0 

Q 

;c 

a 

P 

w 

a 

.. 

-1 

x x x x x x x xxxx XXXI xx 

-- -- . 

x x x x x x x xxxx x xx 

x x x x x x x xxxx xxx xx 

c 

v) 

3 

w 

a 

0 

s 

89 
90 
s IT: 

--5 

-2 

8t 

-a .I 

4 

-a . 

8 

-9 

€ 

8 

I .. S 

t 

i 

L 

91 
4 

z 

P 

a c 

2 

W 

z 

3 

a 

c 

tn’ 

3 

W 

a 

3 

tn 

tn 

‘3W &a 

WQ 

aw 

nM 34 

LO 

2 

a 

m 

A 

U 

92 
I 

W 

m 

3 

ww 

mm 

33 

W 

m 

3 

9 

Y 

0 

U 

m 

93 
Scanivalves inside the model. There are six 48-port, Type S 

Scanivalves in the model, two in the aft fuselage and four in the 

forward fuselage. The nozzle boattail instrumentation will be 

routed from the nozzle, through the inner wing, and to the aft 

Scanivalves, as shown in Figure 6-5. Due to space limitations, 

the pressures will be split between left and right hand nozzles. 

The remaining model static pressures will be routed to the forward Scanivalves. The total precsures for nozzle pressure ratio 

and CMAPS/jet effects high pressure air measurement will be 

routed to a NASA-supplied transducer pod located on the tunnel 

support system. The CMAPS instrumentation required for health 

monitoring and control will also be routed to this transducer 

Pod* 

94 
8 

1 

0 

A 

r 

1 l 

e c 

c 

L 

A 

m 

95 

ORIGINAL PAGZ IS 

OF POOR Q'JALITY 
7.0 ERROR ANALYSIS 

An error analysis was initiated in Phase 1 to assess the random error in key aerodynamic performance parameters (CD, CL, CM) 

as neasured by the model system. Purpose of this activity is to 

identify the major contributing error sources, and any necessary 

changes to the data acquisition/reduction procedures, instrumentation, calibration procedures, and possibly the model design. In 

Phase 1 activity, the error analysis methodology was established, 

the necessary data reduction equations were formulated, and 

errors were predicted for a sample case. This analysis will be 

finalized in Phase I1 after tests of the model under the ANC 

program are completed. This will permit actual model data (i.e. 

measurand values and calibration accuracies) to be used, thus 

ensuring maximum benefit from the analysis. -- - 

7.1 METHODOLOGY - The error analysis procedure being used is a 

comprehensive method developed under the Exhaust System Interaction Program, Reference 14. It has been implemented at MCAIR 

in a generalized error analysis computer program. This program 

was written according to the standard technique of propagation of 

errors as described in Reference 15. The procedure can be described briefly by letting x1# X2r x3 ... represent random values 

whose true values are Xl, X2, X3 ... and letting u represent a 

derived quantity whose true value is given by# 

- 

and also letting el, e2, e3 .. . represent the statistically 

independent errors (standard deviations) of Xi0 X2, x3 ... respectively. Then the error induced in u# which will be denoted 

by 6, as a result of the errors el, e2, e3 ... has a variance 

equal to 

96 
I 

where ;f/axi is the sensitivity of u wit3 respect to the variable 

xi. These partial derivatives are computed by means of either a 

three point or a five point differentiation scheme at values 

equal to the measured values of XI, X20 X3 ... From equation 

(a) 8 the percent contributions to a2 (6) by x2, x3 can be 

computed by 

2 % Contribution = C a (s) /a2 (6) 1 loo. 

ei 

The basic structure of the program is outlined in Figure 7-1. 

The aerodynamic coefficients of interest are calculated in 

several user provided function subroutines. These subroutines 

contain the data reduction equations which are included herein as 

Appendix C. User numerical input consists of the measurands and 

their associated one sigma (standard deviation, a) uncertainties. 

A description of the measurands is also given in Appendix C. 

7.2 SAMPLE CASE RESULTS - A sample test case has been run for 

the simulator test mode to demonstrate the content and fonnat of 

the forthcoming error analysis. The sample case incorporates 

input data fraa tests in Phase 1 of the ANC program, and estimates for the CMAPS related quantities not measured in the ANC 

program. The ANC data, although not expected to accurately represent the current model, does provide a reasonably good data base 

from which the operation and usefulness of the procedure may be 

verified. The numerical results presented here should be interpreted accordingly. 

97 
/ 

MCAIR ERROR ANALYSIS PROGRAM 

i 

/ ,I1 

FUNCTION SUBROUTINES 

TO COMPUTE CL, Cg, CM 

Y 

INPUT DATA 

- MEASUREDVALUES I ‘ I II JJ I - ONESIGMAVALUES 

QPl3480%14 

FGURE 7-1 

ERROR ANALYSIS COMPUTER PROGRAM STRUCTURE 

98 
The operating point selected for the savple case is at Mach / 

0.9, 4' angle of attack, and dry power. This corresponds to a 

typical subsonic cruise condition. The output of the XCAIR error 

analysis program for the sample test case is shown in r'igure 7-2. 

Presented for each calculated quantity (i.e. CL, CD and Cy) are 

its value, variance and standard deviation (sigma). Also listed 

are the percent contributions to the total error of the major 

contributors. Only those measurands contributing more than .1% 

of the total variance are listed in the output. Taking the drag 

coefficient (CD) as an example, the standard deviation is 0.0011, 

or 11 counts, with the balance normal force contributing 68% of 

the total variance and axial force contributing 24%. 

7.3 PLANNED ERROR ANALYSIS CASES - Plans are to complete the 

error analysis in Phase 11, using data from the ANC model to be 

used in the NASA program. Each of the three test modes will be 

analyzed at three Amirks* 9.ese pcfnts havz Seen soiecteci to be 

near important conditions on the DLI mission, as follows: 

DLI MISSION SEGMENT CONDITION 

Supersonic Dash Out Mach 1.4, a = O', A/B Power 

Subsonic Cruise Back Mach 0.9, a = 4', Dry Power 

Subsonic Maneuver Mach 0.6, a = 2O', A/B Power 

99 
SAMPLE CASE OUTPUT 

.CbLC'JLATEr' OU&N'ITIES.o.. JET f FFECTS TEST CASE PT o OAOE in.. 

I PARAMETER 

CL 

co 

CM 

VALUE 

VAR 

SIGMA 

SAMPLE ST OEV 

CONTRlBUTl ON 

PERCENT OF VARIANCE 

MEAS QUANTITIES 

ALPHAM 

A2 

I CPCAV 

FAGA 

FNGA 

M 

MMGA 

PO 

lTR 

EXPLA NATlO N OF TERMS 

0 ESCRl FTlO N 

LIFTCOEFFICIENT, CALCULATED QUANTITY 

DRAG COEFFICIENT, CALCULATED QUANTITY 

PITCHING MOMENT COEFFICIENT, CALCULATED QUANTITY 

VALUEOF CALCULATED OR MEASURED QUANTITY 

VARIANCE OF CALCULATED IlUANTlTY 

STANDARD DEVIATION OF CALCULATED QUANTITY 

STANDARD DEVIATION OF MEASURED QUANTITY 

CONTRIBUTION OF MEASURED QUANTITY TO TOTAL VARIANCE 

PERCENT CONTRIBUTION TO TOTAL VARIANCE 

MEASURANOS 

COMPRESSOR FACE AREA 

AVERAGE CAVITY PRESSURE 

BALANCE AXIAL FORCE READING 

BALANCE NORMAL FORCE READING 

TUNNEL MACH NUMBER 

BALANCE PITCHING MOMENT READING 

TUNNEL STATIC PRESSURE 

TUNNEL TOTAL TEMPERATURE 

MOOEL ANGLE-OF-ATTACK 

Note: Balance errors are taken to be 0.5% of the full scale reading 

OPl~15 

FIGURE 7-2 

ERROR ANALYSIS OUTPUT FOR SAMPLE CASE 

100 
8 0 REFERENCES 

1. Air Force Contract F33615-77-C-2097, "Compact Yulti-Mission 

Propulsion Simul ator Development Program" , 

2. Air Force Contract F33615-77-C-3094, "Advanced Exhaust 

Nozzle System Concepts Demonstration (Part B) 'I, September 

1977. 

3. MCAIR Letter M2Y-343-6710, "Phase I Preliminary Model 

Drawings and Stress Report Data Transmittal", Contract NAS 

2-10791, dated 24 April 1981. 

4. McDonnell Aircraft Company, "Propulsion and Airframe Aerodynamic Interactions of Supersonic V/STOL Configurations" 

(Technical Proposal), MDC A6219, dated 10 December 1979. 

5. Hiley, P. E., Kitmiller, D. E., and Willard, C. N. , "Experimental Evaluation of Non-Axisymetric Exhaust Nozzles", 

AFFDL-TR-78-185, December 1978, Air Force Contract No. 

F3 36 15-76-C-30 19. 

6. McDonnell Aircraft Company, "Turbine Engine Multi-Mission 

Propulsion Simulator Wind Tunnel Demonstration", AFAPL-TR76-73, dated November 1976. 

7. McDonnell Aircraft Company, "Assessment of Installed Inlet 

Forces and Inlet/Airframe Interactions", AFFDL-TR-76-62, 

July 1976. 

8. McDonnell Aircraft Company, "Investigation of Support System 

Interference on a Cylindrical Body", MDC A2188, 26 February 

1973. 

101 
9. K. Lotter, and J. Malefakis, "Intake Design and Intake/ 

Airframe Integration For a Post-Stall Fighter Aircraft 

Concept", Paper No. 32, AGARD Syn'tpOSlUm on High -Angle of 

Attack Aerodynamics, Sandef jord, Xorway, 4-6 October 1978. 

10. Abbott, J. M., "Aerodynamic Performance of Scarf Inlets", 

NASA Technical Memorandum 79055, Lewis Research Center. 

11. General Electric Company, "Final Pretest Report for Propulsion Shulator SN004/1 Aero/Mechanical Test and Evaluation", 

G.E. Technical Memorandum No. 80-210, dated 6 May, 1980. 

12 . McDonnell Aircraft Company, "Experimental Investigation of 

Several Fixed Geometry Inlets for Supersonic Aircraft at 

Mach Numbers from 0 to 2.0", MDC A3566, 17 October 1975. 

13. AEDC, "Evaluation of Boattail Geometry and Exhaust Plume 

Temperature Effects on Nozzle Afterbody Drag at Transonic 

Mach Numbers", AEDC TR-76-102, October 1976. 

14. Glaser, E'. C. and Anderson, R. D., "Performance Projection 

of the Propulsion System Integrated with the Vehicle, Phase 

I1 Exhaust System Interaction Program", McDonnell Aircraft 

Company Report MDC A2817, 12 April 1974. 

15. Mandel, J., "Statistical Analysis of Experimental Data", 

John Wiley h Sons, 1967. 

102 
APPENDIX A 

TURBINE DRIVE AIR SYSTEM PRESSURE LOSS ANALYSIS 

A detailed pressure loss analysis for the turbine drive system was conducted during Phase I of this program. The detailed 

results of this analysis are contained in this Appendix. The 

drive system was divided into the individual components shown in 

Figure A-1. A detailed geometric description of each component 

is contained in Figure A-2, along with the loss factors used and 

flow properties (total pressure, Mach number, and dynamic 

pressure) at each component. 

10 3 
ORIGINAL PAGE IS 

OF POOR QUALITY 

i-- -- -_ 

\ 

\ 

Ii :I1 

“7- 

I. r 

, I 

,I 

....- - 

104 
7 

c 

I- F u 

I 

0 

ORIGINAL PAGE CS 

OF POOR QUALITY ' '' 

1 

0 0d 00 00 0 'I 

I 0 

0 

- 

X.. 

a 

2 

W 

W 

m 

a 

E 

0 0 

Q 

10 5 
APPEXDIX B ._ 

TURBINE BLEED AIR SYSTM D,RESSURE LOSS ANALYSIS 

A detailed pressure loss analysis for the turbine bleed sysI 

I 

tan was conducted during Phase I of this program. The detailed 

results of this analysis are contained in this Appendix. The 

bleed system was divided into the individual components shown in 

Figure B-1. A detailed geometric description of each component 

is contained in Figure B-2, along with the loss factors used and 

flow properties (total pressure, Mach number, and dynamic 

pressure) at each component. 

106 
ORIGINAL PAGE FS 

&, POOR QUALITY 

I3 ’\ 

107 
I 

I 

L 

L., L I/ 

-- 

I 

I 

I 

i 

I 

I 

-1 

10 8 
J 

0 

2 

w 

m 

w 

a 

E 

0 

0 

al 

109 
u) u3 v) 

N 1- 0 01- a 8 

00000 

w- 

> 

4 

a 

;;= 

Ea 

z 

L 

0 

- 0 

c 0 

c V 

L 

LL 

.- 

.- 

- J 8 

n 

w 

w 

m 

c 

110 
APPENDIX C 

ERROR ANALYSIS DATA REDUCTION EQUATIONS 

Presented in this appendix are the preliminary data reduction equations used to calculate CL, CD and CM in the error 

analysis program for the simulator mode. The equations however 

are identical for all three test modes except for the duct 

related parameters. These differences are noted. Also presented 

is a description of the measurands used in the data reduction 

equations. 

The final data reduction equations will be derived as part 

of Phase I1 activity. 

111 
Measurand 

MFSLT 

ACAVN 

-4FASC 

-ASP 

AFB 

AFBU 

AFBL 

AFSW 

AFSWIB 

AFSWOB 

ALPEiAM 

ALP" 

ANFSLT 

APMSLT 

ACAVA 

AWET1 

AWET2 

Description 

Aircraft balance axial force installation 

c~rrection, LB. 

Projected normal cavity area 

Aircraft balance axial force strut cavity 

pressure correction, LB. 

Aircraft balance axial force seal pressure 

correction, LE 

Total pressure-area integrated boattail axial 

force, LB. 

Upper surface pressure-area integrated nozzle 

boattail axial force, LB. 

Lower surface pressure-area integrated nozzle 

boattail axial force, LB. 

Total pressure-area integrated sidewall axial 

force, LB. 

Inboard pressure-area integrated sidewall axial 

force, LB. 

outboard pressure-area integrated sidewall axial 

force, LB. 

Nozzle area assigned to pressure tap I in the 

axial direction. 

Nozzle area assigned to pressure tap I in the 

normal direction. 

Model angle of attack. 

Nozzle angle of attack, DEG. 

Aircraft balance normal force installation 

correction, IN-LB. 

Aircraft balance pitching moment installation 

correction, IN-LB. 

Projected axial cavity area. 

Nacelle wetted area, IN2. 

Nacelle plus nozzle wetted area, IN2 

112 
Measurand 

A2 

CAE 

CABL 

CAFA 

CASW 

I 

I 

CASWIB 

CASWOB 

I 

I CBAR 

CDA 

I CDARON 

CDARONF 

CDAROT 

CDAROTF 

CDB I 

CDBL 

CDBU 

CDNSF 

Description 

Compressor face duct area, IN2. 

Total pressure-area integrated boattail axial 

force coefficient. 

Upper surface pressure-area integrated nozzle 

boattail axial force coefficient. 

Lower surface pressure-area integrated nozzle 

boattail axial 2orce coefficient. 

Aircraft balance corrected axial force 

coefficient . Pressure-area integrated sidewall axial force 

coefficient. 

Inboard pressure-area integrated sidewall axial 

force coefficient. 

Outboard pressure-area integrated sidewall axial 

force coefficient. 

Mean aerodynamic chord. 

Aircraft balance corrected drag coefficient. 

Total nozzle pressure-area integrated drag 

coefficient. 

Total nozzle pressure-area integrated drag 

coefficient plus nozzle skin friction 

coefficient . Total model drag coefficient using pressure-area 

integrated nozzle drag. 

Total model drag coefficient using pressure-area 

integrated nozzle drag plus nozzle skin friction. 

Total pressure-area integrated boattail drag 

coefficient. 

Lower surface pressure-area integrated boattail 

drag coefficient. 

Upper surface pressure-area integrated boattail 

drag coefficient. 

Xozzle skin friction coefficient. 

113 
Measurand 

CDSW 

CDSWIB 

CDSWOB 

CLAROT 

CLB 

CLBU 

CLBL 

CLSW 

CLSWIB 

CLSWOB 

CMARON 

CMAROT 

CMLA 

Description 

Total pressure-araa integrated sidewall drag 

coefficient. 

Inboard pressure-area integrated sidewall drag 

coefficient. 

Outboard pressure-area integrated sidewall drag 

coefficient. 

Nacelle skin friction coefficient. 

Nacelle plus nozzle skin friction coefficient. 

Aircraft balance corrected lift coefficient . (CLAP 

Total nozzle pressure-area integrated lift 

coefficient. 

Total model lift coefficient using pressure-area 

integrated nozzle lift. 

Total pressure-area integrated boattail lift 

coefficient. 

Upper surface pressure-area integrated boattail 

lift coefficient. 

Lower surface pressure-area integrated boattail 

lift coefficient . Total pressure-area integrated sidewall lift 

coefficient. 

Inboard pressure-area integrated sidewall lift 

coefficient. 

Outboard pressure-area integrated sidewall lift 

Coefficient. 

Total nozzle pressure-area integrated pitching 

moment. 

Total model pitching moment coefficient using 

pressure-area integrated nozzle pitching moment. 

Aircraft balance rolling moment coefficient 

transferred to Model Reference. 

114 
Measurand 

CMMA 

CMNA 

CMMAS 

CNB 

CNBU 

. CNBL 

CNFA 

msw 

CNSWIB 

CXSWOB 

CP(1) 

CPCAV 

CPMB 

OCFA 

DDELC 

DEW 

DELCL 

DELCR 

DELNOZ 

FAA 

Description 

Aircraft Salance pitching noment coefficient 

transferred to Model Reference. 

Aircraft balance yawing moment coefficient 

transferred to %ode1 Reference. 

Same as CMMA 

Total pressure-area integrated boattail normal 

force coefficient. 

Upper surface pressure-area integrated nozzle 

boattail normal force coefficient . Lower surface pressure-area integrated nozzle 

boattail normal force coefficient. 

Aircraft balance corrected normal force 

coefficient. 

Pressure-area integrated sidewall normal force 

ccsfffzfent. 

Inboard pressure-area integrated sidewall normal 

force coefficient. 

Outboard pressure-area integrated sidewall 

normal force coefficient . Pressure coefficient at tap I. 

Average cavity pressure coefficient. 

Boattail pitching moment coefficient. 

Aircraft balance corrected side force coefficient. 

Left hand minus right hand canard deflection 

angle, DEG. 

Nozzle deflection due to applied loads, DEG. 

Left hand canard deflection angle. 

Right hand canard deflection angle. 

Nozzle geometric deflection angle, DEG. 

Aircraft balance corrected axial force, LB. 

115 
Mea s ur and 

F.AA2P 

FSA 

ETJAZP 

FRAM 

FSEALA 

FSEALN 

FYA 

NA2P 

F2SEAL 

~~~1~2~3 

KST28-35 

M 

. M2 

t4LA2P 

I MLAPP 

MLAe 

MMAP 

MMA2P 

MMAPP 

MNA2P 

MNAP 

Description 

Aircraft balance axial force corrected for weight 

tare, LB. 

Aircraft balance corrected normal force, LB. 

Aircraft balance normal force corrected for 

weight tare, LB. 

Ram drag, LB. 

Inlet duct seal axial force tare. 

Inlet duct seal nom1 force tare. 

Same as FYGA, LB. 

Same as FYGA, LB. 

Compressor face stream thrust. 

Loads induced model deflection constants. 

Installation effects matrix. 

Tunnel Mach Bumber. 

Compressore face Mach Number. 

Aircraft balance rolling moment corrected for 

weight tare, IB-LB. 

Aircraft balance rolling moment transferred to 

Model Reference. 

Same as MLA2P, IN-LB. 

Aircraft balance corrected pitching moment, 

IN-LB 

Aircraft balance pitching moment corrected for 

weight tare, IN-LB. 

Aircraft balance pitching moment transferred to 

Model Reference. 

Aircraft balance yawing moment correct for weight 

tare, IN-LB. 

Same as MNUP, IN-LB. 

116 
Measurand 

.%VAPID 

MSEAL 

xmsc 

NFB 

NFBU 

NFBL 

NFSW 

NFSWIB 

NFSWOB 

NSF 

P 

PMASC 

PMASCA 

PMASP 

PMB 

PMSW 

PO 

PT 

PT3 

Description 

Aircraft balance yawing moment transferred to 

Model Reference. 

Inlet duct seal pitching moment tare. 

Aircraft balance normal force strut cavity 

pressure correction, LB. 

Total pressure-area integrated boattail normal 

force, LB. 

Upper surface pressure-area integrated nozzle 

boattail normal force, LB. 

Lower surface pressure-area integrated nozzle 

boattail normal force, LB. 

Total pressure-area integrated sidewall normal 

force, LB. 

Inboard pressure-area integrated sidewall normal 

force, LB. 

Outboard pressure-area integrated sidewall 

normal 'force, LB. 

Nozzle skin friction, LB. 

Static pressure, PSFA 

Aircraft balance normal force contribution to 

pitching moment, strut cavity pressure 

correction, IN-LB . Aircraft balance axial force contribution to 

pitching moment, strut cavity pressure 

correction, IN-LB. 

Aircraft balance axial force contribution to 

pitching moment, seal pressure correction, 

IX-LB . Nozzle boattail pitching moment, IN-LB. 

Sidewall pitching moment, IN-LB. 

Static pressure, PSIA. 

Total pressure, PSFA. 

Total pressure, PSIA. 

117 
Measurand 

PTJLF 

PTJLFI (I 

PT2I (I 1 

P2C 

Q 

QO 

RE 

RPM 

SPAN 

sw 

TARE7 

TARE8 

TT 

TTR 

vo 

WAF 

x(1) 

x(2 1 

XCAV 

XMTA 

XNTA 

ZCAV 

ZTA 

Description 

Average nozzle total pressure, PSIA. 

Nozzle total pressure at probe I. 

Compressor face total pressure at probe I. 

Compressor face static pressure, PSIA. 

Tunnel dynamic pressure, PSF. 

Dynamic pressure, PSI. 

Tunnel Reynolds Number. 

Water line of model reference. 

Water line of pressure tap I. 

Fuselage station of pressure tap I. 

Fuselage station of model reference. 

Wing span. 

Wing reference area. 

Internal duct drag correction, LB. 

Body axis ram drag, LB. 

Tunnel total temperature, OF. 

'Tunnel total temperature, OR. 

Freestream velocity, FT/SEC. 

Nozzle mass flow, LB/SEC. 

Nacelle length, IN. 

Nacelle plus nozzle length, IN. 

Moment arm for normal cavity force. 

Axial distance from balance center to model 

reference. 

Lateral distance from balance center to model 

reference. 

Moment arm for axial cavity force. 

Vertical distance from balance center to model 

reference . 

118 
ro 

3 

a 

d 

I4 

E 

nl 

a 

Fi 

PI 

0 

E 

a4 

E 

L E 

I3 0 0 a E baa 

al 

t 

1 

4J 

a 

L 

d) 

k 

m 

m 

Q) 

k 

nl 

a 9) k 

m 

a 

Q) 

k 

m 

PP 

Q) 

k 

nl 

d 

a 

4J 

a 

PI 

- 

m 

0) 

k 

01 

d 

\ 

II 

2 

a 

0 

a 

E 

4 i 

U nl 

.d 

4J 

m 

4J 

cn 

PI 

0 

E 

sr 

Q) 

m 

Q) 

k 

1 

m 

m 

al 

k 

& 

d 

a II 

0 

a 

n 

a n 

a n 

U n 

a n 

Q) 

n 

w A 

m n a n 

P A 

0 

d 

d 

Fi 

d 

119 
v1 

U 

c3 

w 

C 

33 

ww 

nn 

.i 

!i 

3 

E 

0) 

E. a 

% 

a 

m 

3 

3 

a 

Q) 

+, 

8 d 

I 

\o 

al 

k 

1 

P 

d 

kl 

G kc 

8.2 

3 + 

P 

J 

w 

EI 

I 

3 

m 

Q) 

rl 

-6 

a 

E 

3 

m 

4 

d 

d 

Q) 

B 

m 

Q) 

k Q) 1 rl u 

a m 

Q) 

3 

m 

A 

Q) 

k 

a 

CT 

E 

rl 

3 $8 -a 

U 

d 

Q) 

a 

O 

r, 

rl n 

P n 

P n 

a 

c3 

rl 

d 

d 

In 

d 

120 
n 

E- H !-I CI- Y 

a 

ka 

I 

mH 

a- 

&& 

&U 

..I.. 

n 

H 

Y 

3 

4 

n 

H 

Y 

& 

U 

W 

0 

II 

a 

2 

cr 

z 

n 

d 

n 

H 

Y 

3 

4 

n 

H 

Y 

0, 

U 

W 

0 

II 

4 

z 

a 

E 

A 

P 

n 

H 

Y 

3 

n 

H 

Y 

& 

W 

W 

0 

II 

a 

4 

h 

4 

n 

a 

h 

3 

a 

I 

n 

H 

Y 

P 

a 

a 

I 

II 

Y 

n 

H 

4 

p1 

a 

Y 

5 

n 

w 

n 

n 

H 

rl 

Y 

z 

I: 

& 

a 

n 

U 

h 

H 

Y 

U 2 

n 

H 

a 

U 

W 

Y 

0 

II 

r: 

b 

a 

m 

n 

cn 

n 

n 

H 

Y 

ii & 

U 2 

a 

n 

n 

H 

U 

Y 

n 

H 

& 

U 

Y 

W 

0 

+ 

a 

4 

a 

E 

Z 

+ 

E 

2 

II 

lk 

2 

m 

n 

F 

4 

kl 

4 

+ 

3 

E 

4 

II 

Q 

m 

2 

n 

.d 

0 

cu 

121 
?D 

J 

I 

. 

n 

n 

H 

rl 

z 

a 

n 

Y 

E 

U 

n 

H 

Y 

5 

4 

U 

n 

H 

LL 

U 

Y 

W 

0 

11 

3 

co 

z 

LL 

a 

n 

al 

n 

n 

c( 

Y 

3 LL 

a 

n 

n 

U 

U 

Y 

5 

4 

U 

n 

H 

a4 

U 

Y 

W 

0 

+ 

a 

E 

O 

-4 

U 

4 

k 

U 

E 

.rl 

F 

8 

k 

4 

Q) 

!4 

3 

m 

m 

al 

k nl 

rl 

rl 

m 

E 

2 

m 

aim 

dQ) 

NU 

N& 

00 

zw 

0 

m 

m 

m 

Q) 

k 

Q) n k n 

5 a 

m 

m 

Q)nnnnW s n 

n 

Y 

s 

4 

A 

H 

Y 

nl 

U 

W 

0 

I. 

II 

H 

3 

M 

a 

a 

2 

n 

0 

n 

H 

Y 

s 

4 

n 

Y 

Y 

LL 

u 

w 

0 

I 

II 

m 

a 

B 

5 

n 

a 

n 

H 

Y 

5 

8 

4 

n 

n 

Y 

W 

0 

II 

m 

H 

3 

v1 

E 

Z 

a 

n 

m 

n 

U 

Y 

3 

4 

A 

U 

LL 

U 

Y 

W 

0 

II 

a 

B v1 

Lc 

z 

n 

P 

a 

si 

M 

E 

z 

+ 

H 

3 

m 

E 

2 

II 

m 

5 

E 

z 

n 

vc 

a 

8 

2 

;13 

+ 

m 

2 

2 

II 

E 

n 

P 

U 

\ 

n 

U 

\ 

U 

\ 

n 

U 

\ 

a 

4 n W GI 

E 

z 

a 

Y 

n 

d 

Y 5+ + 

s 

4 

U 

3 

z 

‘3 

hl 

II 

(u 

II 

(u 

It 

122 
mnuanapuaalvc 

o-rooooooooo 

d-d-rQ-rd-aWaWW 

...a*...... 

n 

n 

5 

Y 

n 

a 

Lo 

U 

3 

cn 

Y 

U 

\ 

L5 

s 

Y 

cy 

I1 

m 

H 

s 

cn 

Z 

W 

n 

d 

n 

3: 

cn 

n 

Y 

n 

Y 

a 

U 

\ 

n 

cn i! 

k! 

Z 

cy 

Y 

n 

m 

!3 

8 

n 

n 

n 

n 

?i 

Y 

n 

a 

m 

Y 

U 

\ 

n 

I4 

5 

Y 5 

cy 

n 

E3 

3 

cn 

4 

u 

H 

n 

U 

n 

X 

v) 

n 

Y 

n 

a 

m 

Y 

U 

n 

\ 

8 

Y 3 

cy 

n 

P 

cn 

B 

3 

n 

a 

P 

; 

+ 

m 

U 

L 

cn 

3 

It 

5 

8 

A 

vc 

al 

& 

1 

m 

W 

al 

& 

a 

0 

Q 

123 

n 

3 

u 

I 

2 

Y 

n 

3 

Y 9 

8 

U 

n 

?! 

2 

Y 9 

u 

I 

Y 

n 

8 

U 

A 

3 u 

E 

v 3 

Y 

+ 

A 

E 

d 

m 

n 

c;r 

4 

u 

+ 

m 

Y 

n 

3 

Y 4 

E 

d 

W 

n 

cl 

u 

+ 

3 

P 

W 

It 

LII 

u 

8 

m 

m 

n 

w 
m 

LI 

.rl 

C s 

3 n 

9) 

U 

I OdOOOOOOOOO 0 ......... e. 

cn ln#4mFLnLneFFFF 

k m~u mpapua ow Q) Q)ww Q) Q) u IUP oa 

00000000000 

aFaFem*aaaa 

e.......... 

d 

mzmmmmmmm 

... .- . 

t4 

1 II++ 

OD 

m nnnn 

a 

m 

z 

CI 

u z 

c3 

u 

+ 

z 

4 

c1 

u 

oa + 

m 

Y 

+ + 

c3 

u 

I1 

z 

m 

3 W 

+ m 

E 

u 

a 

tl 

u s v1 

P 

u It 

a 

z 

4 

U 

2 

n 

It z 

4 

u 

2 

n 

z 

2 

2 

u 

s 

cn 

u 

n 

d 

NQ) 

Ok n n n n n 

NIU 

zm m a u a Q) CCI 

n 

a n 

P n 

U n 

a 

0 

P 

0 

a 

124 
I 

E 

3 

9) 

u 

k 

0 

tl] 

a 

0 

Q) 

k 

& 

8 

OD 

"f 

d 

8 

m uaaw ot3oooo .. ... 

n 

2 

0 

4 

u 

m 

a 

F 

W 

n 

0 

4 

r) 

8 

id 

a 

F 

W 

1 

F 

W 

n 

OD 

4 

8 

4J 

(d 

n 

: 

0 

4 

AJ iu 

a 

v 

W 

n 

OD 

4 

OE 

125 
Q) 

U m P gal0 Bal P u w 

0 0 . . .dZ k malo I rlrl. 

0 9 -0 00 000 -00 00 -000 

f13 rlrlrl rlrl drlrlrl~rldrlrldrlrl 

00 oooQ)cnq~qgoao ... 

al 

?I 

a 

0 

. 

0 

+ 

4 

n n 

P U 

n 

cv 

4 

- 

- 

E 

* 

a 

4 P 

E 

II 

cv 

I: 

Y 

n 

a 

.. 

a 

5 

n 

Q) 

126 

a 

n 

YI 
m 

J 

Q) 

U 

i( 

1 

0 

tll 

8 

4 

4 

W 

-d 

a 

Q) 

m 

3 

U 

3 

n 

3 

Q) 

d 

c 

W 

rl 

O 

. 

r( 

a 

Q) 

E 

d 

w 

al 

a 

z 

0 

JJ 

It 

k 

% 

E 

rl 

w 

Q) 

a 

'p, 

P 

0 

U 

It 

73 

E 

4 

VI 

Q) 

a 

0 

4 

It 

n 

a n 

P n 

0 n 

a 

n 

n 

3 

Y 3 

v1 

0 

U 

U 

5 ttl 

I 

(Y 

tc 

II 

E! 

4 E 

n 

U 

127 
n 

!i 

U 

4 

0 

Y 

n 

a 

Y 

n 

3s 

U 

01 

0 

I 

II 

Y 

z: 

Irr 

4 

2 

n 

m 

n 

4 

u 

4 

0 

% 

Y 

n 

a 

% 

x 

Y 

n 

0 

II 

U 

m 

Y 

i!! 

n 

0 

n 

% 

W 

N 

n 

0 

m 

Y 

$ 

7 

Y 

II 

4 

U 

3 al 

n 

a 

m 

J 

n 

n 

n 

H 

Y 

d 

rl 

a: 

n 

U 

n 

H 

Y nl 

U 

U 

W 

0 

II 

n4 

m 

4 

Y 

a 

$ 

n 

m 

U 

n 

H 

n 

Y 

4 

rl 

4 

n 

U 

n 

H 

Y 

6 

U 

Y 

w 

0 

I 

II 

cc 

a 

3 LL 

n 

P 

a n 

L n 

el 

d 

II 

H 

. 

r- 

\o 

d 

N 

0- 

zm 

0 

‘CT 

4 

. 

+ 

3 

rl 

Y 

I 

n 

el 

Y 

n 

cv 

Y 

II 

11 

tc 

m 

W 

R 

nn 

PU 

128 
01 

u 

.rl 

c 

3 

m . . . . . . . . . ... 

129 
P 

(Y 

4 

z 

I: 

II 

2 Z 

E 

h 

a, 

P 

3 

3 

I: 

II 

E 

n 

w 

130 
I 

dl .. 

r( 

Q) 

a 

0 

E 

0 

e, 

h 

9) 

W 

I 

E 

h 

e, 

E 

tu 

U' 

d 

E 

9) 

0 

c 

4 

d 

4 

P 

Q) 

e,U 

WUE 

dQ) 

h& 

UQ) 

hw 

4Q) rda 

0 

fi 

4 

n 

i! 

Y 

n 

4 

& 

2 

X 

I 

01 

I: 

II 

nl 

P 

Y 

s 

P 

n 

. 

3 

Q 

m 

4J 

E 

9) 

rl 

0 

rl 

W 

W 

P) 

8 

m 

rl 

X 

m 

'8" 

P 

Q1 

UP) 

EO rdc 

dQ) 

dL 

PQ w 

49) 

wcr; 

Lld 

UP) 

Lla 

40 

RI: 

0 

m 

. 

!2 

nnn 

nnn 

131 
rll 

J 

E 

4 

m 

n 

4 Fr 

Y 

8 

+ 

A 

3 

eL 

a 

8 

Y 

m 

n 

4 

83 -= 

n 

P 

m 

Fr 

2 

2 z 

8 4 

P 3 U 

W 

+ 

II 

VI 

+ + 

4 

W 

II 

E-r 

4 

U 

n 

2 

n 

4 Q 

W 

II 

E-c 

W s 

II 

E-c 

4 

E 

U 

2 

n 

m n 

P A 

U 

h 

a A 

Q) 

132 
I 

I 

! 

I 

I 

I 

I 

I 

i 

i 

1 

I 

1 

I 

~ 

i 

I 

, 

1. Repat No. 

NASA CR-177369 

2. Gomnmmt Accuion No. 3. Rwpimt’t CIulog No. 

PROPULSION AND AIRFRAME AERODYNAMIC INTERACTIONS SeDt 85 

OF SUPERSONIC V/STOL CONFIGURATIONS 6. Performing Orpnizrtion codr 

4. Title md Subtitle 5. Reoon Date 

PHASE I - FINAL-REPORT 

7. Author(r1 

14. Sponsoring Agency codc 

505-43-01 

12. Sponwring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546 

8. Performing Orpnization RIport No 

I 

5. Supplemtrry Nota 

M. R. Mraz and P. E. Hiley , 

9. hrformtng orgmniutton Nrme and Addrrrr 

McDonnell Aircraft Company 

P.O. Box 516 

St. Louis, Missouri 63166 

Point of Contact: 

(41 5) 694-6265 Ames Research Center, Moffett Field, CA 94035 

Technical Monitor, Rodney 0. Bailey 

10. Work Unit No. 

T- 3 288Y 

11. Contract or Grant No. 

NAS2-10791 

13. Type of Repor! nd hid Covered 

6. Abstract 

9 kcuritv asoif. (of this report) 

Unclassified 

A wind tunnel model of a supersonic V/STOL fighter configuration has been 

tested to measure the aerodynamic interaction effects which can result 

from geometrically close-coupled propulsion systemlairframe components. 

The approach was to configure the model to represent two different test 

techniques. One was a conventional test technique composed of two test 

modes. 

measured, including inlet/airframe interactions. In the Jet-Effects mode, 

incremental nozzle/airframe interactions are measured. The other test 

technique is a propulsion simulator approach, where a sub-scale, externally 

powered engine is mounted in the model. This allows proper measurement of 

inlet/airframe and nozzl e/airframe interactions simultaneously. 

In the Flow-Through mode, absolute configuration aerodynamics are 

20. Security CIasuf. (of this pmgc) 21 No of P~QM 22 Rice' 

Uncl ass i f i ed 132 

7. Key Words (Suggcrted by Author(st1 1 18. Distribution Statement 

Flowfield Interactions 

Propulsion Simulators 

Supersonic V/STOL Star Category 02