AD-A009 138 DESIGN AND CALIBRATION OF THE ARL MACH …

AD-A009 138

DESIGN AND CALIBRATION OF THE ARL MACH 3 HIGH REYNOLDS NUMBER FACILITY

A. W. Fiore, et al

Aerospace Research Laboratories Wright-Patterson Air Force Base, Ohio

January 1975

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I REPORT NUMBER

ARL 75-0012 2 GOVT ACCESSION NO 3 RECIPIENT'S CATALOG NUMBER

Al)-/) Co 7 iM 4. TITLE (and Submit)

DESIGN AND CALIBRATION OF THE ARL MACH 3 HIGH REYNOLDS NUMBER FACILITY

5. TYPEOF REPpRT A PERIOD COVERED Final Technical July 1965 - December 1972 6. PERFORMING ORG. REPORT NUMBER

7. AUTHORfi; ». CONTRACT OR GRANT NUMBERfsJ

A. W. Fiore; D. G. Moore; D. H. Murray and J. E. West, Maj. USAF

9 PERFORMING ORGANIZATION NAME AND ADDRESS

Fluid Mechanics Research Laboratory and Theoretical Aerodynamics Rsch Lab (ARL/LF & LH) Wright-Patterson AFB, OH 45433

10. PROGRAM ELEMENT, PROJECT, TASK 4RFl.a WDRKJIWIT NUMBERS

61102F; 7065-14-28 61102F; 7064-02-06

11, CONTROLLING OFFICE NAME AND ADURESS

Aerospace Research Laboratories (AFSC) Bldg. 450, Area B Wright-Patterson AFB, OH 45433

12. REPORT DATE

January 1975 13. NUMBER OF PAGES

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Wind Tunnel Mach Number 3 High Reynolds Number

Design Calibration

PRICES SUBJECT TO CHANGE 20. ABSTRACT (Continue on ravers« aide It necessary mnd Identify by block number)

The design philosophy and details of the ARL Mach 3 High Reynolds Number Facility are described and the results of extensive calibration tests are presented. The facility is an intermittent cold flow wind tunnel having a constant area test section 8 inches wide and 8.2 inches high, with a test rhombus 23 inches long. It operates in a blowdown mode using dry compressed air which is exhausted to the atmosphere through a silencer. Tunnel stagnation pressures range from 40 psia to 570 psia, with stagnation temperatures in a range from 400 degrees Rankine to 500 degrees Rankine depending upon ambient conditions. The corresponding free

DD FORM 1 JAN 71 ]473 ED,'r,rtfcl ^P ' WAV/ ««i \K nR<;ni FTF

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20. (Cont'd)

stream unit Reynolds numbers vary from 7 million per foot to 140 million per foot The tunnel has been operated with a vacuum exhaust which permitted stagnation pressures as low as 4.2 psia with a corresponding free stream unit Reynolds number of 0.8 million per foot. The aerodynamic calibration measurements were made at nominal stagnation pressures of 100, 300 and 500 psia with an average stagnation temperature of 480 degrees Rankine. Test rhombus determinations included lateral and longitudinal Mach number distributions and flow angularity measurements. A limited number of tunnel blockage tests were performed and some ,'iow visualization tests were made to determine characteristics of the start-stop process.

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PREFACE

This report was prepared jointly by the Fluid Mechanics Research

Laboratory of ARL under Project 7065, "Aerospace Simulation Techniques,"

and by the Hypersonic Research Laboratory of ARL under Project 7064, "High

Velocity Fluid Mechanics." This report covers the design and calibration of

the ARL Mach 3 High Reynolds Number Facility.

The authors wish to acknowledge the significant contributions of

Mr. Emil Walk, formerly of the Fluid Mechanics Research Laboratory, as the

project monitor, and Ms. Constance M. Woehle and Mrs. Shirley E. McGrath

for typing the manuscript.

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•TABLE OF CONTENTS

SECTION PAGE

I INTRODUCTION 1

II DETAILED WIND TUNNEL DESIGN . 4

1. HIGH PRESSURE AIR SYSTEM 5

a. Storage Tanks 6

b. Manifolding 6

c. Piping 6

2. WIND TUNNEL PROPER 7

a. Tunnel Foundation 7

b. Pressure Control System 8

c. Settling Chamber 9

d. Entrance Bellmouth 12

e. Nozzle 12

f. Test Section 17

g. Model Support Systems 19

h. Diffuser 20

3. EXHAUST SYSTEM 22

a. Vacuum 23

b. Atmospheric 24

III WIND TUNNEL CONTROLS AND INSTRUMENTATION 25

1. SYSTEM STATUS MONITORS 26

2. INTERLOCK SYSTEM 28

3. STAGNATION PRESSURE CONTROL SYSTEM 30

a. Process 30

b. Pressure Sensor 31

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TABLE OF CONTENTS (CONT'D)

SECTION PAGE

c. Programmer 31

d. Controller 31

e. Servo Amplifiers 32

f. Process Control Loop 32

4. VALVE RESPONSE TESTS 33

5. WATER HAMMER 33

6. CONTROLLER RESPONSE 35

7. MODEL SUPPORT SYSTEM CONTROL 35

IV AERODYNAMIC CALIBRATION OF THE WIND TUNNEL 37

1. THE LATERAL MACH NUMBER DISTRIBUTIONS 37

2. THE LONGITUDINAL MACH NUMBER DISTRIBUTIONS 43

3. THE TUNNEL EMPTY MACH NUMBER DISTRIBUTION 45

4. THE CENTERLINE RMS MACH NUMBER 46

5. FLOW ANGULARITY MEASUREMENTS 46

6. BLOCKAGE TESTS 52

7. REMARKS ON FLOW LIMITATIONS 53

V CONCLUSIONS 56

REFERENCES 60

APPENDIX: BASIC DESIGN CONSIDERATIONS 131

1. INPUT DESIGN PARAMETERS 131

a. Test Section Mach Number 131

b. Desired Model Sizes 132

c. Desired Reynolds Number Range 132

d. Desired Run Times 132

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TABLE OF CONTENTS (CONT'D)

SECTION PAGE

2. CALCULATED PARAMETERS 133

a. Tunnel Stagnation Conditions 133

b. Tunnel Pressure Ratio, A 135

c. Test Section Size 136

d. Mass Flow 137

e. Exhaust System 138

f. Air Supply 141

3. OUTPUT DESIGN PARAMETERS 142

a. Test Section Mach Number 143

b. Test Section Size 143

c. Stagnation Temperature 144

d. Stagnation Pressure 145

e. Run Time 145

f. Exhaust System 146

g. Performance Envelope 146

LIST OF SYMBOLS 169

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LIST OF ILLUSTRATIONS

FIGURE PAGE

1 Perspective Drawing of the ARL Mach 3 High Reynolds Number Facility 62

2 Estimated Pressure Drop in Distribution Piping as a Function of Mass Flow and Storage Pressure 63

3 Mach 3 Nozzle Configuration 64

4 Original Exhaust System Configuration-Vacuum 65

5 Rise of Sphere Pressure With Run Time for a Range of Stagnation Pressures 66

6 Present Exhaust System Configuration-Atmospheric 67

7 Mach 3 Process Diagram 68

8 Logic Symbol Convention 69

9 Control Valve Status 69

10 Isolation Valve Control Logic 70

11 Run Interlock Logic 71

12 Exhaust Valve Control Logic 71

13 Alarm System Logic 72

14 Stagnation Pressure Control System 73

15 Mach 3 Process Control Loop 74

16 Valve Response Test-2 inch Valve 75

17 Valve Response Test-6 inch Valve 76

18 Controller Test-2 inch Valve 77

19 Controller Test-6 inch Valve 78

20 Scale Drawing of the Mach 3 Nozzle and Test Section 79

21 Scale Drawing of the Mach Number Survey Rake 80

22 Survey Rake Reference Axis and Nomenclature 81

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LIST OF ILLUSTRATIONS (CONT'D)

FIGURE PAGE

23a Impact Mach Number Versus y/As for xp/xr = - 0.690 at p0 = 109 psia and T0 = 430oR 82

23b Impact Mach Number Versus V/AS for xp/xr = - 0.345 at p0 = 111 psia and T0 = 477°R 83

23c Impact Mach Number Versus y/As for xp/xr = 0 at p0 = 111 psia and T0 = 4840R 84

23d Impact Mach Number Versus y/As for xp/xr = + 0.345 at p0 = 108 psia and T0 = 4770R 85

23e Impact Mach Number Versus y/As for xp/xr = + 0.690 at p0 = 103 psia and T0 = 4720R 86

24a Impact Mach Number Versus y/As for xp/xr = - 0.690 at p0 = 306 psia and T0 = 4680R 87

24b Impact Mach Number Versus y/As for xp/xr = - 0.345 at p0 = 316 psia and T0 = 4820R 88

24c Impact Mach Number Versus y/As for xp/xr = 0 at p0 = 323 psia and T0 = 491

0R 89

24d Impact Mach Number Versus y/As for xp/xr = + 0.345 at p0 = 313 psia and T0 = 483^ 90

24e Impact Mach Number Versus y/As for xr = + 0.690 at p0 = 318 psia and TQ = 476°R 91

25a Impact Mach Number Versus y/As for xp/xr = - 0.690 at Po = 518 psia and T0 = 4780R 92

25b Impact Mach Number Versus y/As for xp/xr = - 0.345 at p0 = 520 psia and T0 = 4910R 93

25c Impact Mach Number Versus y/As for xp/xr = 0 at p = 532 psia . and T0 = 499

0R 94

25d Impact Mach Number Versus y/As for xp/xr = +0.345 at p0 = 517 psia and T0 = 491;bR 95

0.690 at 25e Impact Mach Number Versus y/As for xp/xr = + p0 = 523 psia and T0 = 486lbR 96

26a Centerline Mach Number Versus Longitudinal Distance for Po = 100 psia, T0 = 4680R, where Re/ = 20.69 x 106 per foot . . 97

vn

.

FIGURE

26b

26c

27a

27b

27c

28a

28b

28c

29a

29b

29c

30

31

32

33

34a

34b


Centerline Mach Number versus Longitudinal Distance for Po = 300 psia, T0 = 480.06oR, where Re/H = 59.2 x 106 per foot .

Centerline Mach Number versus Longitudinal Distance for Po = 500 psia, To = 489.05oR, where Re/£ = 95.0 x 106 per foot .

Mach Number versus Longitudinal Distance for r/As = ± 1.0 at p0 = 100 psia, T0 = 468.40R and Re/£ = 20.69 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 1.0 at Po = 300 psia. To = 480.06oR and Re/l = 59.2 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 1.0 at Po = 500 psia, T0 = 489.05oR and Re/Z = 95,0 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 2.0 at Po ~ 100 psia, To = 468.40R and Re/Z = 20.69 x ID6 per foot

Mach Number versus Longitudinal Distance for r/As = ± 2.0 at Po = 300 psia. To = 480.06^ and Re/Z = 59.2 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 2.0 at Po = 500 psia, T0 = 489.05oR and Re/£ = 95.0 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 3.0 at Po = 100 psia. To = 468.40R and Re/£ = 20.69 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 3.0 at Po = 300 psia. To = 480.06oR and Re/£ = 59.2 x 106 per foot

Mach Number versus Longitudinal Distance for r/As = ± 3.0 at Po = 500 psia. To = 489.05oR and Re/£ = 95.0 x 106 per foot

Wall Mach Number versus Longitudinal Distance for Tunnel Empty Condition with Reynolds Number as a Parameter ....

Root-Mean=Square Mach Number versus Unit Reynolds Number Per Foot

Scale Drawing of Flow Angularity Wedge

Scale Drawing of Flow Angularity Cone

Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = - 0.690 for po = 103.6 psia with Settling Chamber Spreader Cone Tip Downstream ....

Normalized Surface Presswre Difference versus Angle of Attack at xp/xr = - 0.34b for po = 100.6 psia with Settling Chamber Spreader Cone Tip Downstream ....

PAGE

98

99

100

101

102

103

104

105

106

107

108

109

110

in

112

113

114

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FIGURE PAGE

34c Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = 0 for po = 100.6 psia with Settling Chamber Spreader Cone Tip Downstream 115

34d Normalized Surface Pressure Difference versus Angle of Attack at xp/xr » 0.345 for p0 = 95.0 psia with Settling Chamber Spreader Cone Tip Downstream 116

34e Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = 0.690 for p0 = 95.4 psia with Stagnation Section Spreader Cone Tip Downstream 117

35 Velocity Profile in a Model Settling Chamber with and without a Spreader Cone . 118

36 Flow Angularity versus Lateral Distance at xp/xr = 0 (Rhombus Center) with the Spreader Cone Tip Upstream .... 119

37 Flow Angularity versus Lateral Distance at xp/xr = 0 (Rhombus Center) with the Spreader Cone Tip Downstream ... 120

38 Unit Reynolds Number Effect on the Flow Angularity at xp/xr - 0 (Rhombus Center) 121

39 Flow Angularity versus Lateral Distance with Longitudinal Distance as a Parameter for pg = 100 psia, T0 = 470oR, M = 2.99 and Re/£ = 20.6 x 10g per Foot 122

40 Flow Angularity versus Longitudinal Distance at z/(b/2) = + 0.5 (East Side) with the Spreader Cone Tip Downstream . . 123

41 Flow Angularity versus Longitudinal Distance at 2/(b/2) = 0 (Centerline) with the Spreader Cone Tip Downstream 124

42 Flow Angularity versus Longitudinal Distance at z/(b/2) = - 0.5 (West Side) with the Spreader Cone Tip Downstream . . 125

43 Flow Angularity versus Probe Roll Angle at the Center of the Test Rhombus (xp/xr = 0) 126

44 Tunnel Blockage versus Mach Number 127

45 Schlieren and Shadowgraphs of Flow Over a Two-Dimensional Wedge at M = 3.0 and Re/£ = 20.66 x 106 per Foot for a = 0° and a = 4° 128

46 Schlieren and Shadowgraphs of Flow Over a Two-Dimensional Wedge at M = 3.0 and Re/£ = 20.66 x 106 per Foot for a = 7° and a = 10° 129

IX

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APPENDIX

FIGURE PAGE

A-l Design Flow Chart 150-153

A-2 Streamwise Location of Transition on a Model at Mach 3 as a Function of Free Stream Unit Reynolds Number 154

A-3 Effect of Air Condensation on Tunnel Operating Range at a Mach Number of 3 155

A-4 Rate of Change of Free Stream Unit Reynolds Number with Tunnel Stagnation Temperature at Mach 3 156

A-5 Variation of Free Stream Unit Reynolds Number with Tunnel Stagnation Pressure and Temperature 157

A-6 Tunnel Pressure Ratio Requirements I58

A-7 Variation of Maximum Model Size with Fineness Ratio for Shock Reflection Limited Cases 159

A-8 Variation of Maximum Base Diameter for Starting Tunnel with Blunt Cone Models 160

A-9 Variation of Maximum Cone Length with Inviscid Test Section Size for Shock Reflection Limited Cases 161

A-10 Variation of Maximum Wedge Length with Inviscid Test Section Size for Blockage Limited Cases 162

A-ll Mass Flow Per Unit Test Section Area as a Function of Tunnel Stagnation Conditions 163

A-12 Variation of Silencer Size with Test Section Size for a Silencer Intake Velocity of 100 ft/sec 164

A-13 Variation of Required Vacuum Volume with Test Section Size for a 60 Second Run at a Unit Reynolds Number of 1.2 x 106 per foot. . 165

A-14 Variation of Air Storage Volume Requirements with Test Section Size and Operating Pressure 166

A-15 Maximum Model Reynolds Number Attainable as a Function of Inviscid Test Area 167

A-16 Predicted Tunnel Performance Envelope 168

LIST OF TABLES

TABLE PAGE

I Mach 3 Nozzle Coordinates 15

A.I Turbulent Flow Testing Capabilities 148

A,II Laminar Flow Testing Capabilities with an Atmospheric Exhaust . 149

xi

SECTION I

INTRODUCTION

In July 1965 the Aerospace Research Laboratories at Wright-Patterson Air

Force Base began to give serious consideration to the construction of a high

Reynolds number aerodynamic test facility, the intention being to study

turbulent flow phenomena at supersonic and low hypersonic speeds. The need

for aerodynamic test data at high Reynolds numbers, particularly in this

speed range, had become increasingly apparent in the course of identifying

critical problem areas associated with the fluid mechanics of new and advanced

Air Force systems. Increased attention was being given to "low and fast"

aircraft and other weapon delivery systems, and there was also a trend toward

larger rocket boosters and reusable launch vehicles. Fundamental information,

such as heat transfer data, was scant in the appropriate Mach number/Reynolds

number range, and extrapolation was virtually impossible. It was clear that

there was a need for experimental data pertaining to pure turbulent boundary

layers, flow interactions, heat transfer rates, skin friction values, wake

characteristics, and aerodynamic stability under turbulent flow conditions.

The Mach number range of particular interest appeared to extend from Mach 2 Q

to Mach 6, WK.,i Reynolds numbers ranging up tO'10 or greater. A thorough

review of existing supersonic and hypersonic facilities throughout the

country^ " ' revealed that very few were designed for operation at free

stream unit Reynolds numbers much above 10 per foot. The principal exceptions

were the shock tunnels which are unsuitable for some types of detailed flow

studies due to their extremely short running times. In view of the very

limited high Reynolds number testing capability which the country had, there

appeared to be little doubt that the establishment of a high Reynolds number

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Simulation capability at ARL would be a sound and timely investment in a

research and development tool necessary for servicing current and future needs

of Air Force systems.

Convinced of the need, ARL undertook to study the problem of providing

the desired simulation capability as quickly and as inexpensively as possible,

using existing technology wherever possible. It was apparent early in the

study that the Mach number range could not be satisfactorily covered with one

flow channel. The logical approach appeared to be to have a facility con-

sisting of an unheated leg for low Mach numbers and a heated leg for higher

Mach numbers, with as much common equipment as possible. It was decided that

representative Mach numbers for the two legs should be Mach 3 and Mach 6,

with no immediate provisions to be made tor varying the test section Mach

number of each leg.

Based upon the above decisions, a preliminary design study was made in

which the requirements for major components were analyzed to establish the

overall technical feasibility of the approach, and to ascertain costs, man-

power requirements, and the effect of the facility on the existing research

complex. Test section sizes and run times were of prime consideration due

to their direct impact on facility size and facility service demands. Follow-

ing this study, which concerned itself with conventional blowdown wind tunnel

designs, two intensive reviews and several small studies were made to see if

new or emerging techrology could be expected to produce significantly superior

facilities which would warrant a delay in the construction of a new facility.

Since this did not appear to be the case, in February 1967 the decision was

made to proceed with the design and construction of the Mach 3 leg of the

facility. In-house work on the detailed design started immediately and

construction effected by the FluiDyne Engineering Corporation of Minneapolis,

Minnesota, was completed approximately two years later. This was followed by

an exhaustive series of component check-out tests comprising well over one

hundred tunnel runs. These were completed during March 1969. The installation

and check-out of research instrumentation occupied the facility until

September 1969, when calibration tests were started. These were completed in

July 1970 and regular aerodynamic testing began during August 1970. The

first aerodynamic test report containing data obtained from the tunnel was

published in October 1971.'7'

The purpose of this report is to provide a single published source of

information on the facility. It is intended to be useful to those preparing

for tests in the facility and to those interpreting data obtained in the

facility. For the interest of those engaged in the design and construction

of similar facilities, an appendix covering the basic design considerations

which led to the detailed design described in Section II of the report has

been included.

SECTION II

DETAILED WIND TUNNEL DESIGN

The detailed design of the wind tunnel was based upon the design

criteria and calculations discussed in the appendix. A facility design goal

was established, namely, the safe production and control of a uniform Mach 3 o

flow, having a free stream unit Reynolds number (IteJ of 10 per foot at a oo

stagnation pressure (P ) of 570 psia and a stagnation temperature (T ) of

500oR. The test section size was to be nominally 8 inches square, with

allowances for nozzle boundary layer growth. The tunnel was to be designed

for direct exhaust to the atmosphere, and for exhaust to an existing 3

100,000 ft vacuum sphere capable of being evacuated to a pressure of 1 torr.

Based upon this design goal, a facility concept was developed, and expanded

to the point of specifying design or performance criteria for all significant

systems, subsystems and components. The end result of this effort is depicted

in the facility perspective presented in Figure 1.

In order to identify the major components in the system, a brief

description of the facility is given in terms of the path taken by the air

through the facility. Air passes from a high pressure storage area through

two parallel four-inch diameter Schedule XXS high pressura pipes into two

remotely controlled isolation valves, which are manifolded together on the

downstream side. This manifold connects to the common intake of a six-inch

diameter pressure control valve and a two-inch diameter pressure control

valve, of which both are hydraulically actuated. The two valves are connected

in parallel to permit individual selection for operation at either high or

low mass flow. Air passes from the active control valve to the settling

chamber through a wide angle expansion, which incorporates a flow spreader to

promote a uniform velocity profile. The settling chamber experiences the

full stagnation pressure of the flow, and is fitted with a rupture disc to

protect it from overpressurization in the event of control valve failure.

Screens in the settling chamber help to reduce the scale and intensity of the

turbulence in the flow prior to acceleration through the aerodynamic nozzle to

the test section Mach number of 3. After the air passes over the model in the

test section, it undergoes some deceleration in the diffuser, and then enters

the downstream ducting and exhaust system. Further details on the facility

components are given below with the same flow sequence.

1. HIGH PRESSURE AIR SYSTEM 3

The ARL high pressure air system includes over 15,000 ft of 3000 psi

air storage for the use of many special purpose test rigs and wind tunnels.

Three four-stage reciprocating air compressors are able to take atmospheric

air and deliver it to the storage vessels at a combined mass flow rate of

1.0 lbm/sec. In addition to removal of condensed water and oil by separators,

oil vapor is removed by a special "oil-sorb" unit. The saturated air is

then dried by passage through one of two automatically cycled silica gel

drying towers, which permit dew points as low as -100oF to be achieved. Fine

mesh filters assure that the air passing to the storage tanks is essentially

particulate free. One 8250 ft section of the storage area is normally

available to the facility. The maximum system pressure is 3000 psia, but the

pressure available to the Mach 3 facility is often lower due to the operation

of other facilities. The air consumption/recovery ratio can be as high as

200 to 1 for this facility, which means that a sequence of three 60 second

high pressure runs consumes more air than can be recovered in ten hours of

compressor plant operation. Further details of the high pressure air system

of interest to this particular facility are given below.

5

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a. Storage Tanks 3

The 8250 ft of storage volume referred to above Is made up of

3 3 twenty-eight 250 ft tanks and one 1250 ft tank. The tanks contain no heat

sink materials and are completely exposed to atmospheric conditions. The

small tanks each have a 2 in.-2500# ASA Ring Joint Flange outlet connection.

The corresponding internal diameters are 1.503 inches and 2.300 inches,

respectively, which are sufficient to keep local outlet velocities below

50 ft/sec for the small tanks and 100 ft/sec for the large tank.

b. Manifolding

The storage tanks are connected to two four-inch diameter Schedule

XXS pipe manifolds through individual high pressure gate valves. Each pipe 3

manifold is connected to fourteen 250 ft tanks, but only one is connected 3

to the 1250 ft tank. The manifolds therefore have connected volumes of 3 3

3500 ft and 4750 ft . Under normal conditions corresponding manifold air

velocities should not exceed 150 ft/sec and 200 ft/sec, respectively.

c. Piping

The manifolds are connected to the facility by independent

four-inch diameter Schedule XXS pipes, each incorporating a pressure balance

four-inch high pressure shut-off gate valve upstream of the corresponding

facility isolation valve. With the usual allowances for elbows, valves, tees

and reducers, the approximate equivalent lengths of the two supply lines,

including manifolds, are 400 ft and 300 ft, the longer length being associated 3

with the manifold connected to the 1250 ft tank. Figure 2 shows how the

pressure drop increases with the mass flow rate for the combined supply lines

for a range of storage pressures. Losses have been estimated up to, but

T

excluding, the control valves. These curves can be used to establish the

minimum storage pressure (PST) required for a run, providing allowances are

made for the pressure drop through the control valve, and for the pressure

drop in the storage vessels due to the air consumed and the associated exp

sion cooling. From an examination of the wide range of operating conditions

possible, a good rule-of-thumb would be to have a minimum storage pressure of

at least thêe times the tunnel stagnation pressure.

All of the components of the high pressure air system described

above wore subjected to a hydrostatic test at 4500 psig prior to use.

2. WIND TUNNEL PROPER

The wind tunnel proper comprises those mechanical elements essential to

the conveyance, control, and aeodynamic conditioning of the test air. The

design of the most important elements is discussed below, primarily from the

standpoint of function and physical characteristics. Control and instrumenta-

tion aspects of the wind tunnel are discussed in Section III.

a. Tunnel Foundation

The tunnel foundation has the function of supporting the tunnel

components, the upstream piping, and a portion of the downstream ducting. It

is designed to resist the horizontal thrust loads that occur at maximum mass

flow conditions, and the vertical loads that would occur following the rupture

of either of two eight-inch rupture discs. By combining the foundations of

the various components into one unit, the opposing horizontal forces are

cancelled out in the foundation, and are not transmitted to the ground. The

foundation is of reinforced concrete containing sleeved anchor bolts for

component mountings. It is vibrationally isolated from the building slab by

a 1/2-inch peripheral expansion joint.

b. Pressure Control System

The high pressure air control system comprises two isolation valves,

one primary control valve, one secondary control valve, one equalization

valve, and a control system. The function of the overall system is to

control discretely the stagnation pressure in the settling chamber during a

run.

(1) Isolation Valves

The two isolation valves are four-inch Grove Series G gate

valves which are provided to assure positive remote isolation of the facility

from the high pressure air supply system. In addition, the electro-hydraulic

actuators operate fast enough (less than one second) for the valves to be used

as shut-off valves under emergency conditions if the pressure control valve

malfunctions in the open position. An adjustable differential pressure

switch across the isolation valves can be used to time-sequence the operation

of the control valves. In the event of power failure, the valves are

hydraulically actuated to the closed position to provide a failsafe condition.

(2) Pressure Control Valves

The primary pressure control valve is a six-inch Annin

Model 4510 valve, with a C.G.S. Model 361 electro-hydraulic actuator. It has

a Cv of 290 and is used to control mass flows over a range of 50 to 200

lb /sec. The secondary pressure control valve is a two-inch Annin Model 4510,

with a C.G.S. Mode? 321 actuator. It has a Cv of 35 and ir used to control

mass flows below 50 lb /sec. Both valves have linear characteristics, and

can be operated in either an automatic mode or manual mode to achieve a set-

point operating pressure in the settling chamber.

(3) Eoualization "alve

The equalization valve is a Jamesbury 3/4-inch Type HP ball

valve which provides a means of pressurizing the upstream side of the control

valves prior to operation of the isolation valves. The equalization valv? is

fitted with a failsafe spring-loaded air cylinder operator which closes upon

air or electrical power failure.

(4) Control Elements

The stagnation pressure control elements consist of a

programmer, two controllers, two servo amplifiers, and three pressure sensing

transducers which are integrated with the control valves and actuators to form

the control loops for establishing and maintaining the desired air flow

conditions. These form a process control system which is described later in

Section III.

c. Settling Chamber

The settling chamber has the function of conditioning the air before

it passes through the nozzle. Conditioning includes promoting a uniform

velocity profile at the nozzle entrance, and reducing the scale and intensity

of turbulence. In addition, the settling chamber volume must be sufficient

to eliminatp time variations in the sensed stagnation pressure, so that any

control valve perturbations are not amplified in the control loop. Particular

features of the settling chamber assembly are discussed below.

(1) Shell Assembly

The settling chamber shell assembly constitutes a pressure

vessel, designed for 700 psi dir service in accordance with Section VIII of

the ASME code for unfired pressure vessels. ^ The upstream end of the

welded structure incorporates a flanged manifold for mounting the two control

^ ■,..■•-*

valves, followed by a 90° long radius elbow. A conical wide-angle expansion

member connects the downstream side of the elbcw to the main cylindrical

member of the shell. This main member incorporates a downstream connection

flange, thrust mount pads, and a rupture disc assembly mounting flange. The

eight-inch diameter rupture disc is sandwiched between flanges, and is suffi-

cient to prevent overpressurization of the vessel in the event of control

valve failure (wide open) at the maximum system pressure of 3000 psi. Thrust

forces are transmitted by the brackets to the tunnel foundation through a

steel thrust support member. Welded to the top of vessel are three mounting

pads for the tunnel siJe-wall swing-arm assembly. The settling chamber

diameter was chosen to provide a mass-averaged velocity of approximately

30 ft/sec. One of the primary concerns in the design was that äir temperatures

would on occasion be considerably below -20oF, where many carbon and low

alloy steels begin to suffer serious decreases in impact resistance. Instead

of using a stainless steel, or of complying with the special material impact

tests requirements, the code option of designing to a pressure equal to

2 1/2 times the maximum working pressure was chosen. The cylindrical section

of the vessel did not require special low temperature consideration due to the

use of an inner steel liner not subject to pressure loads. Since run tim2s

are short, on an absolute time scale, the pressure vessel itself does not

reach the low air flow temperature.

(2) Flow Spreader

The limited building space available to the facility made it

necessary to have a 90° long radius elbow, immediately followed by a wide-

angle diffuser section, in order to make the transition from the high pressure

piping to the full settling chamber diameter. Because of the high rrass flow

10

i^.|l...IV»Tilv.v,n,w.^.w.!l -• --

rates there was considerable concern that turbtlence and asymmetry at the

control valve, and centrifuging in the elbow, would lead to significant flow

nonuniformities in the settling chamber. To counter this possibility it was

decided to incorporate a perforated cone with the wide-angle diffuser. The

idea of using such a device was not new, having been mentioned by Ferri and (g)

Bogdonoff as early as 1954/ ' but no experimental performance data were

available. In fact, there was considerable room for debate on whether such a

device might be more effective pointing downstream, rather than upstream as

depicted in the few literature references discovered/ " ' In view of this,

some small scale studies were made by ARL to determine the influence of

orientation on effectiveness. These studies are discussed further in connec-

tion with flow angularity measurements in Section IV. Pending the outcome of

the small scale tests, the flow spreader was designed to be reversible. The

cone has a semi-vertex angle of 45° (the semi-expansion angle of the wide-angle

diffuser is 30°) and is fabricated from 3/4-inch thick perforated steel plate

having .707 in. diameter holes and a porosity of 36%. The base is reinforced

with a steel ring to support and align the cone inside the shell. Movement

of the spreader is prevented by four steel stops welded to the shell upstream

and by steel spacer rings downstream.

(3) Screens

Three turbulence screens are used downstream of the flow

spreader to reduce the scale and intensity of turbulence and to improve

further the velocity profile across the chamber. The screens can be variously

located within the shell, depending on the cone orientation and ring spacer

positions, but can never b. placed closer than approximately one shell dia-

meter to the cone apex, in order to avoid possible adverse wake effects. They

11

.rvwr^rT^'-.--

are fabricated from stainless steel wire cloth silver soldered to a steel ring.

Fairly high stresses can be experienced in taut mesh screens, so that the

effects of low temperature brittleness, cloth to ring bonding, uniformity of

screen tautness, and shock loading must be considered in estimating allowable

stresses and arriving at acceptable mesh sizes. The effect of a screen failure

can be remarkably detrimental to the'finish of a precision machined nozzle

and model. The Mach 3 facility uses one screen having a mesh size of

16 x 16 x .015 in. and two screens having mesh sizes of 22 x 22 x ,010 inches.

Original screens having mesh sizes of 20 x 20 x .010 in. and 32 x 32 x .0068 in.

failed during check-out testing and were replaced by the heavier gauge ones.

Screen spacings are 167 and 250» expressed in terms of the usual ratio of

screen separation distance to screen wire diameter.

d. Entrance Bellmouth

The entrance bellmouth makes the transition from the circular

cross-section of the settling chamber to the rectangular cross-section of the

nozzle blocks. It consists of top and bottom flat aluminum plates, and

circular aluminum side blocks. The side blocks attach flush to the straight

nozzle sidewalls and extend upstream at a 14-inch radius to the settling

chamber inner diameter. The flat plates are tangent to the upstream ends of

the nozzle blocks (41° 38 ft from horizontal), and also extend to the settling

chamber diameter. The bellmouth blocks are a precision fit with the nozzle

block assembly, and with the bellmouth plates which they carry.

e. Nozzle

The nozzle is the single most important component of the facility,

and as such received considerable attention in the design, fabrication, and

12

......,-.^- .. ■ ■■

■

■

installation processes. The nozzle assembly was designed to permit the use

of different nozzle blocks producing Mach numbers up to 4.5, but to date the

facility has been operated exclusively at Mach 3.0 with the original blocks.

(1) Aerodynamic Design

The inviscid contour of the nozzle is based upon a completely

(131 (14) analytical design method due to Friedrichs^ ' as modified by Nilson.x '

(151 It was successfully used by Baronv ' for the design of a number of supersonic

nozzles at the Naval Supersonic Laboratory of the Massachusetts Institute of

Technology in 1954. The mathematical method of characteristics is not

directly employed; rather a truncated series solution of the nonlinear wave

equation is used to express the pertinent flow properties adjacent to the

nozzle axis. The solution is valid in both the subsonic and supersonic

portions of the field, and no assumptions need to be made with respect to the

disposition of the sonic line. The method computes characteristic lines of

the field by a numerical integration process, and downstream from one of these

the flow may be made uniform by a simple mass-flow criterion for the simple-

wave region streamlines. By examining the series it is possible to estimate

the magnitude of the errors introduced by the discarded terms, whereas a

comparative check using the method of characteristics requires recomputation

with a finer mesh size. Reference 15 contains conveniently tabulated

coordinates for the design characteristics, and corresponding potential-flow

nozzle coordinates, for a Mach number range from 1.5 to 3.5. Direct use was

made of the Mach 3.0 potential-flow data in arriving at that portion of the

contour downstream of the nozzle inflection point. A circular arc profile was

used for the contour upstream of this point. The circular arc was fitted to

the fixed ordinates of the throat and inflection point, and was made tangent

13

- ■

to the slope at the inflection point. It should be noted that the arc was

fitted after correcting the inviscid contour for boundary layer effects. The

boundary layer displacement thickness was calculated from Burke's equation ( '

which relates the local turbulent boundary layer displacement thickness to the

local Mach number and Reynolds number in the following way:

6* M i.sn = 0.0463 ~

Re 0.276 X

The displacement thickness on the nozzle at the exit plane was calculated to

be 0.0231 in. and 0.0482 in. for free-stream unit Reynolds number of 10 per

foot and 7 x 10 per foot, respectively, corresponding to stagnation pressures

of 570 psia and 40 psia at a stagnation temperature of 500oR. To avoid the

possibility of recomprcssion in the nozzle at low Reynolds numbers, the

largest viscous correction in the operating range must be used. The value of

0.0482 in. was rounded off to 0.0500 in., and was used for both the sidewall

correction and the contoured wall correction. Rather than calculate 6* as a

function of x down the nozzle, the displacement thickness was assumed to be

a linear function of x, starting from zero at the nozzle throat. The

magnitude of the correction is such that any inaccuracy so introduced is

negligible. In order to preserve the benefits of plane parallel sidewalls,

the total viscous correction was applied to the contoured walls. The nominal

test section size of 8.0 in. high by 8.0 in. wide therefore became an actual

size of 8.2 in. high by 8.0 in. wide. Figure 3 summarizes the nozzle

configuration, and Table I gives the nozzle coordinates.

(2) Mechanical Design

The nozzle assembly is made up of top and bottom contoured

aluminum nozzle blocks, two steel nozzle block supports, and two steel

14

prTjsw«""^"'.''"" "

Table I

MACH 3 NOZZLF COORDINATES

¥ X

The contour of the Mach number 3 nozzle is based on Friedrich's method and is corrected for boundary layer growth. The subsonic intake and the supersonic starting portion to x = 4.6535 (the inflection point in the contour) are formed by a radius of 15.548 inches with the center located on the ordinate through x = 0, the nozzle throat. Since the coordinates are given upstream of the inflection point by this radius, the coordinates given below are for the contour from the inflection point to the nozzle exit.

0. 0.9445 4.6535 1.6572 4.6830 1.6662 4.7971 1.7006 4.9123 1.7352 5.0000 1.7614 5.0285 1.7699 5.1459 1.8046 5.2644 1.8397 5.3841 1.8748 5.5000 1.9087 5.5049 1.9101 5.6268 1.9454 5.7503 1.9810 5.8748 2.0165 6.0000 2.0519 6.0006 2.0521 6.1275 2.0878 6.2558 2.1235 6.3852 2.1592

6.5000 2.1905 6.5158 2.1948 6.6479 2.2306 6.7810 2.2661 6.9157 2.3019 7.0000 2.3240 7.0513 2.3374 7.1884 2.3729 7.3267 2.4084 7.4663 2.4437 7.5000 2.4521 7.6072 2.4789 7.7493 2.5140 7.8924 2.5495 8.0000 2.5749 8.0372 2.5837 8.1838 2.6184 8.3313 2.6530 8.4800 2.6872 8.5000 2.6918

15

r* ,**-» -.-- r-nrn^pw.^

Table I (continued)

8.6301 2.7214 8.7817 2.7553 8.9345 2.7891 9.0000 2.8033 9.0887 2.8225 9.2443 2.8557 9.4012 2.8887 9.5000 2.9092 9.5595 2.9215 9.7193 2.9545 9.8804 2.9861

10.0000 3.0094 10.0430 3.0178 10.2070 3.0497 10.3723 3.0809 10.5000 3.1047 10.5391 3.1120 10.7073 3.1426 10.8771 3.1729 11.0000 3.1945 11.0481 3.2030 11.2207 3.2326 11.3948 3.2617 11.5000 3.2790 11.5704 3.2906 11.7474 3.3192 11.9259 3.3473 12.0000 3.3587 12.1059 3.3750 12.2876 3.4024 12.4706 3.4293 12.5000 3.4335 12.6553 3.4558 12.8414 3.4819 13.0000 3.5038 13.0291 3.5078 13.2186 3.5328 13.4095 3.5575 13.5000 3.5690 13.6021 3.5819 13.7964 3.6057 13.9921 3.6291 14.0000 3.6300 14.1897 3.6521 14.3889 3.6745 14.5000 3.6866 14.5896 3.6963 14.7925 3.7179

14.9968 3.7389 15.0000 3.7421 15.2028 3.7593 15.4106 3.7795 15.5000 3.7877 15.6200 3.7987 15.8316 3.8178 16.0000 3.8323 16.0477 3.8362 16.2598 3.8539 16.4768 3.8711 16.5000 3.8729 16.6956 3.8880 16.9165 3.9041 17.0000 3.9100 17.1393 3.9198 17.3642 3.9348 17.5000 3.9431 17.5909 3.9486 17.8196 3.9631 18.0000 3.9734 18.0506 3.9763 18.2833 3.9889 18.5000 4.0002 18.5181 4.0011 18.7555 4.0124 18.9949 4.0232 19.0000 4.0234 19.2366 4.0333 19.4804 4.0427 19.5000 4.0434 19.7265 4.0517 19.9479 4.0595 20.0000 4.0603 20.2257 4.0672 20.4789 4.0737 20.5000 4.0742 20.7346 4.0797 20.9926 4.0850 21.0000 4.0851 21.2530 4.0898 21.5000 4.0933 21.5162 • 4.0935 21.7819 4.0966 22.0000 4.0984 22.0503 4.0988 22.2801 4.1000

16

■"■^"■c- "pv—r"—^T->

sidewalls. These components are bolted to each other and to the nozzle

flange which mates with the downstream flange on the settling chamber. In

this manner, and with the aid of linear "0"-ring type seals, a pressure tight

box was designed to withstand 700 psig upstream of the throat, 200 psig

downstream of the throat, and a pressure of 1 torr throughout. The nozzle

blocks are aluminum plates eight inches wide by 32-3/8 inches long, and are

bolted and keyed to the steel support members which transmit axial and

vertical loads from the blocks into the nozzle flange. The sidewalls of the

box are steel flat plates which are flanged at the upstream ends to provide

a means of attachment to the nozzle flange. The design pressure of 200 psig

downstream of the nozzle throat relates to the stagnation pressure attainable

behind a normal shock at a Mach number of 3 and a stagnation pressure of

600 psi. Higher pressures due to emergency conditions are not reached due to

the rupture disc located further downstream.

f. Test Section

(1) Aerodynamic Design

A closed test section was chosen in contrast to an open jet,

since it was felt that at the high operating densities the strong shear layer

at the jet boundary might induce significant secondary flow disturbances, and

might also lead to a significant noise problem. A nominally square test

section was chosen in preference to any other cross-sectional shape due to a

wide variation in anticipated test model configurations. Geometrically then,

the test section is a parallel wall continuation of the nozzle exit cross-

section. No further corrections for boundary layer growth were made initially,

for reasons of simplicity in fabrication and the ever present uncertainty of

computed boundary layer displacement thicknesses. It is therefore to be

17

■

■

■^rywT *:rm?77-?^

expected that a slight negative Mach number gradient will exist in the test

section due to the uncompensated boundary layer growth.

(2) Mechanical Design

The test section comprises a top wall, a bottom wall, and two

sidewalls. The top and bottom walls are bolted to the nozzle flange upstream

and to the diffuser plates downstream. In addition to the end support, this

framework is supported from below by a movable nozzle-test section cart which

can be adjusted to achieve vertical and lateral alignment during assembly.

The sidewalls attach to the top and bottom walls with quick-release latches,

and are supported during assembly and removal by a swing arm support

permanently mounted to the settling chamber. Sealing of the various rectan-

gular components comprising the nozzle-test section assembly is accomplished

with linear 0-ring type seals, which include several "tee" intersections

requiring careful handling during assembly. Once assembled, the components

form a pressure-tight box which is designed for 200 psig. The sidewalls

extend from the diffuser inlet to a point ten inches upstream of the nozzle

exit, in order to provide maximum window coverage. Three sets of sidewalls

were fabricated, one set being furnished with eight-inch diameter window

assemblies centered on the nozzle exit plane, the others being blank for future

window locations as desired. All the sidewalls are symmetrical, and can be

turned end for end to double the window location possibilities. The window

assemblies themselves comprise two-inch thick schlieren quality glass discs,

permanently mounted in steel frames to insure a flush fit with the sidewalls

at all times. The top wall of the test section accommodates a four-inch

diameter steel blank which can be replaced by similarly sized instrumentation

plugs as desired. The plan bottom wall originally installed in the test

18

r'.T^"'^^ vW-l»,-<JV-"i'W?7¥T^H',-""-"l~''"

■.itVÛBWUHJ'

section can be replaced by either of two other walls which are associated with

the two model support assemblies described later.

g. Model Support Systems

The tunnel is equipped with two model support assemblies, one

providing a pitch capability of jJO0, and one giving a fixed zero pitch angle.

Both assemblies have their own test section bottom wall to permit complete

bench setup of models and instrumentation. The sting socket on each strut is

identical in design to permit the interchange of stings. The maximum design

loads were +2000 lbs. normal force, acting at a point on the sting centerline

three inches upstream of the plane of intersection of the model base with the

sting, and +1000 lbs. axial force, acting along the sting centerline. Maximum

moments were taken as those resulting from the application of the maximum

normal and axial forces acting either together or independently. Since the

maximum loads indicated result from an abrupt nonsteady flow condition, the

model support systems were designed for an impact factor of two. If required,

the tunnel can be operated in a "clean" condition by employing a plain test

section bottom wall furnished with the tunnel.

(1) Fixed Strut

The fixed strut model support system comprises a straight

strut support mounted on a plate which is attached to the test section bottom

wall. The strut and mounting plate can be removed from the test section as

an assembly without removing the test section bottom wall if desired. The

upper end of the strut contains a tapered socket for mounting the sting. The

back of the socket contains a sting nut which is used for both seating and

unseating the sting plug in the socket. The main length of the strut has a

wedge-shaped leading edge, and an instrumentation cavity machined into the

19

trailing edge to route leads from the model to the outside through openings

machined in the mounting plate. The machined cavity is slotted to accommodate

an insert cover plate, and a screwed cover plate behind the sting socket

completes the mechanical assembly,

(2) Movable Strut

The movable strut model support system comprises a strut,

strut support assembly, actuator, potentiometer, and enclosure. The movable

strut is similar to the fixed strut except that it is a circular arc segment

supported and guided by a support assembly bolted to the test section bottom

wall. A hydraulic cylinder actuator is trunnion-mounted to the strut support

casting, and is attached to the strut with a clevis bracket. The full piston

stroke provides exactly the required +10° pitch angle range, and a linear

potentiometer provides remote readout of the strut pitch angle. Provision has

been made for centering the pitch angle about the tunnel centerline. The

entire lower portion of the assembly, including the actuating cylinder and

potentiometer, is enclosed in a pressure-tight enclosure which is sealed and

bolted to the test section bottom wall. The lower half of the enclosure is

removable to provide access to the inner assembly without disturbing

electrical and hydraulic feedthroughs.

h. Diffuser

The diffuser was designed as a constant area duct, having the same

cross section as the test section, followed by a diverging section. A

diffuser throat configuration was considered, but not employed for the

following reasons. At Mach 3 the theoretical ratio of the diffuser throat

area to the test section area is 0.7192 for a clean tunnel. This value is

based upon the diffuser throat being just sufficient to swallow the test

20

■ ■

section normal shock during starting. However, the presence of a model in the

test section during the starting process creates a shock system with greater

losses, resulting in a lower total head downstream. On the basis of mass flow

continuity the diffuser throat must therefore be larger than the theoretical

clean tunnel value, and in practice a 30% increase in throat area ratio is

usually employed. Applying this factor to the calculated value at Mach 3

raises the diffuser throat area to 0.9350 of the test section area. In view

of the closeness of this ratio to unity, and in view of the thickening

boundary layers on the diffuser walls, there appeared to be little point in

providing a diffuser convergence. However, the requirement was established

that the constant area duct be designed to accommodate top and bottom throat

plates at a later date, if desired. The resulting diffuser subassembly

consists of a constant area duct section, a diverging transition section, a

slip joint, and a safety tee.

(1) Constant Area Section

The constant area section is made up of four separate

machined plates, bolted and keyed together to form a rectangular duct 8.0 in.

wide x 8,2 in, high x 80.0 in. long. The plates are sealed with linear 0-ring

type seals, and can be individually replaced or provided with inner blocks or

plates to provide a different aerodynamic configuration. The duct assembly

is entirely supported by the adjacent components to which it is bolted.

(2) Transition Section

A transition from rectangular to circular flow cross section

is accomplished by a machined weldment which changes from an upstream internal

cross section of 8.0 in. x 8.2 in. to a downstream internal diameter of

13.25 inches. This allows the flow to diverge to more than twice the area

21

^T-».;^*-?,:'! mjKui.. .*vm *w*iw"''-J*'y.-'-- ■■'^"-'"w 's^' - -■■^■.r™,-,.,™i,-.J: rrr

over a length of 38.0 inches. The downstream end of the transition piece is

continued as a straight pipe section for 6.5 in. to form part of the slip

joint which follows. The weldment is supported by a movable cart, which

allows alignment of the diffuser in the same manner as that of the test

section cart, and in turn supports the downstream end of the constant area

duct.

(3) Slip Joint and Safety Tee

This is a dual purpose component designed to provide a slip

joint for removing upstream equipment, and to provide a mounting flange for a

rupture disc assembly. A 14 x 14 x 8 in. weld tee is used to accomplish

this. The upstream end is flanged, and has a machined socket to accept the

downstream end of the transition section to form the slip joint. A movement

of 2.5 in. is possible, which is sufficient to allow removal of the nozzle,

test section, or diffuser duct. An eight-inch diameter rupture disc is

mounted on the top outlet flange of the tee. The disc has a rating of 200

psig and prevents overpressurization of all components upstream as far as the

nozzle throat. The rupture disc assembly is sandwiched between the tee

flange and the lower flange of a vent stack which penetrates the building

roof. Replacement of the disc requires only that the flanges be slightly

separated by jack screws. The downstream flange of the tee connects with

the exhaust ducting, and the tee itself is supported by a thrust mount tied

to the tunnel foundation.

3. EXHAUST SYSTEM

As discussed later in the appendix, the tunnel was initially designed 3

to exhaust into a 100,000 ft vacuum sphere, due to uncertainties about the

availability and effectiveness of low pressure drop silencers, and this

22

puiNii .. ..JI,' !■ ■r.ovqjfw^ Tr«****r*rwr7^r-T" - -

system is described in paragraph a. below. Operational experience with this

system pointed up a significant low frequency sphere vibration problem which

gave rise to concern for the continued structural integrity of the sphere.

Due to ambient and localized temperature effects, appendages, and widely

differing ways in which the sphere was used to serve several facilities, a

meaningful determination of vessel fatigue life was considered unreliable.

It was therefore decided to disconnect the Mach 3 wind tunnel from the

vacuum sphere, and to utilize a common exhaust silencer for it and the Mach 6

wind tunnel, which was subsequently completed in 1972. This atmospheric

exhaust system is now in use and is described in paragraph b. below.

a. Vacuum

The vacuum exhaust system configuration was as depicted in

Figure 4, The entire system, including the vacuum sphere, was designed for

service between 60 psig pressure and 1 torr pressure. The wind tunnel was

isolated from the vacuum sphere by the 36 in. diameter sphere valve. The

only purpose of the 14 inch diameter tunnel valve was to provide personnel

safety in the event of sphere valve failure with an open test section. The

vent valve permitted depressurization of the system to atmospheric pressure

if required. The 90° turn was designed as a tee to provide a full 36 in.

access hatch to the model catcher, which was installed at 45° to the incident

flow. Apart from the sphere vibration noted above, the system performed

entirely satisfactorily. The system performance of sphere pressure versus

run time is given for the full range of tunnel stagnation pressures in

Figure 5. Assumptions made include an initial sphere pump-down to 1 torr,

and no vacuum pumps on-line during tunnel operation. In practice the vent

valve was automatically operated when the sphere pressure reached 45 psig

23

to insure against inadvertent rupture of the sphere safety disc, which was

designed to fail at 60 psig.

b. Atmospheric

The currently used atmospheric exhaust system is depicted in

Figure 6. Much of the vacuum exhaust system was used to construct the new

one, particularly the ducting, sphere valve, and model catcher tee. The

tunnel valve was eliminated and replaced by ring spacers, and the vent valve

was replaced by a rupture disc designed to fail at 20 psig. The silencer

assembly consists of two identical commercial units mounted on a concrete

plenum which contains a guide vane assembly. The commerical units are

rectangular panel insert types which employ a glass fiber acoustic fill

material between perforated steel plates. The units have good dynamic

insertion loss ratings at frequencies between 850 and 3400 cycles per

second, but measurements to date indicate that much of the acoustic energy is

at low frequencies outside of this range. The units are operated with a

maximum face velocity of 4500 fpm, compared to a rated maximum of 5000 fpm.

Certified performance data on this type of unit indicates that, even when the

face velocity approaches zero, the overall attenuation does not improve

significantly. It is possible that acoustical lagging of the ducting and

silencer will be necessary, since the total radiated noise at a distance of

50 ft from the silencer exhaust was reduced only from 126 dB with one silencer

unit to 113 dB with a second unit stacked immediately on top of the first.

Efforts to positively identify the major noise sources and frequencies are

continuing.

24

ipWt^Wf -.I'lfV .\m*»W!-!*r

SECTION III

WIND TUNNEL CONTROLS AND INSTRUMENTATION

The instrumentation required to monitor the condition or status of the

tunnel at all times, the interlocks which insure safe operation, the stagnation

pressure controller, and the model support control system will be described in

this section. A prime consideration throughout is to provide a maximum of

safety to personnel and equipment while employing the simplest and most

trouble free control hardware. The controls and interlocks originally used

for the "exhaust to sphere" configuration have been modified, or removed as

required, for the "exhaust to atmosphere" configuration. Consideration will

be limited to the latter configuration and the safety interlock with the

parallel Mach 6 facility.

Figure 7 is a simplified process diagram illustrating all of the

necessary monitoring and control stations. With all valves initially closed,

a typical run sequence would be as follows:

1) Preset the desired P0 on Controller

2) Open "Supply" valves

3) Open "Exhaust" valve

4) Open "Isolation" valves

5) Initiate "Run" utilizing proper "Control" valve for mass flow

dvsired

6) Close "Control" valve

7) Close "Isolation" valves

8) Close all remaining valves.

25

i]l»,W.Î,.T''--"J^''^^,",^'--*,^''''J|'"""^'-'<-'-*",""?''fa''^^w*'J*WJ"

1, SYSTEM STATUS MONITORS

The flow variables are monitored and indicated to the operator by the

following devices (see Figure 7):

1) P^, Pp—Supply Pressure

Ashcroft pneurratic transmitter (C4080TA)

Ashcroft pneumatic receiver (1224C)

0-3000 psia calibration

2) P0--Settling Chamber Pressure

Ashcroft pneumatic transmitter (C4480S)

Ashcroft pneumatic receiver (1228)

0-600 psia calibration

3) T --Settling Chamber Temperature

Conax copper constantan thermocouple probe (T-SS12-B-PJFC-PG2- 125AT-18)

Assembly Products panel meter (355)

-200 to 100oF calibration

4) P --Test Section Static Pressure—dual range

Range 1: Taylor pneumatic transmitter (215TA11112-1507)


0-1000 mm Hg calibration

Range 2: Ashcroft pneumatic transmitter (C4G30S)


30 inch vacuum to 300 psig calibration

5) T --Test Section Static Temperature (Wall)

Copper constantan thermocouple

Assembly Products panel meter (355)

-200 to 100oF calibration

26

Tf <-*-*■-'-"■-■- -■'---. ^,... _-.. T^*%mww**wityJ¥*im*'

6) P --Exhaust Pressure

Ashcroft pneumatic trarvimitter (C4080S)


30 inch vacuum to 60 psig

In addition, certain flow parameters are sensed by pressure switches and

used by the interlock system. They are as follows (see Figure 7):

1) AP--Differential pressure across isolation valves

Deltadyne MElOl-B-A-R-ll

Set to 15 psid

2) P --Settling Chamber Pressure

Mercoid Type DA-21-2

Set to 600 psia

3) APSRr) --Settling Chamber Rupture Disk Limit

Deltadyne ME102-E-A-R-21

Set to 1/2 psid

4) APrjnp—Diffuser Rupture Disk Limit

Deltadyne ME102-E-A-R-21

Set to 1/2 psid

5) Pe--Exhaust Pressure—Dual Range

Barksdale D2T-A80

Set at 20 psia

To complete the interlock system inputs, each valve is fitted with a

limit switch at each end of its stroke, thereby giving four logical states to

each valve motion:

1) closed

2) not closed

27

j'twwm- .....

3) not open

4) open

States 2 and 3 are used to indicate a valve in some position between fully

closed and fully open.

2. INTERLOCK SYSTEM

The interlock system utilizes 24 volt DC relay and limit-switch logic

elements throughout to insure reliable service and easy maintainability. All

final control elements are 24 volt solenoid valves which control either air or

hydraulic actuators as described in Section II of this report. Figures 9

through 13 illustrate the detailed logic involved in each of the valve opera-

tions. Figure 8 illustrates the "and," "or" logic convention. Figure 9

illustrates the "and" logic circuit which gives the combined status of the

Control valves. Figure 10 illustrates the complete logic diagram for the

Equalization and Isolation valves. The normal operating sequence is first to

preselect either Isolation valve #1 or #2 or both and then to initiate an

"OPEN" cycle which opens the Equalization valve until the AP across the

Isolation valves drops below 15 psid, and then the preselected Isolation

valve(s) are opened. From the logic diagram it can be seen that, in order to

actuate the Equalization valve, eight inputs are required to be "true"

simultaneously:

1) the "OPEN" pushbutton must be actuated,

2) the hydraulic pressure must exceed the high limit set point,

3) the Mach 6 air supply valve must be closed,

4) the cabin doors must be closed,

5) the control valves must be closed,

6) the "CLOSE" pushbutton must not be actuated,

28

7) the exhaust valve must be open, and

8) the isolation valves must not be open.

As soon as the Equalization valve signal is output, the "OPEN" pushbutton may

be released and the signal will be maintained. Similarly, the Isolation

valve signal will be output when:

8

9

Once the I so

1

2

3

4

5

the "CLOSE" pushbutton is not actuated,

the exhaust valve is open,

the settling chamber pressure is less than 600 psia,

the hydraulic pressure is greater than the low limit set point,

the equalization valve signal is "true,"

the control valves are closed,

the delta pressure across the isolation valve is less than 15 psid,

the calibration valve is closed, and

one or both of the isolation valves have been selected.

ation valve signal has been output it will remain as long as:

the isolation valve remains open,

the "CLOSE" pushbutton has not been actuated,

the exhaust valve remains open,

the settling chamber pressure does not exceed 600 psia, and

the hydraulic pressure does not drop below the low limit set point.

As soon as the Isolation valve(s) open, the Equalization valve will close,

since the "Isolalion Valves Not Open" signal will be lost.

Each of the remaining logic diagrams. Figures 11 through 13, similarly

indicate the interlock conditions required for each of the other valves and

the "ALARM" circuit.

29

LtHWIT" 'i'' '.

3. STAGNATION PRESSURE CONTROL SYSTEM

The stagnation pressure control is a closed loop electro-hydraulic servo

system with major components connected as illustrated in Figure 14. The

control system will increase the settling chamber pressure from the initial

starting value to any preselected set-point pressure between 20 and 570 psia

within three seconds and maintain this pressure to within ±0.5% or +0.5 psia,

whichever is larger, for a maximum run time of 60 seconds. It may be operated

in either manual or automatic mode and in automatic mode is interlocked and

initiated as outlined in the preceding discussion of interlocks. It is

normally operated in the automatic mode with manual control being used for

maintenance checkout purposes only.

a. Process

The process is assumed to be influenced only by the settling

chamber and nozzle, and it can be shown that the transfer function is of the

form:

GT(s) = ^(s) / K(s) = ^

where PQ is t+ie settling chamber pressure in psia,

rfi is the pressure control valve mass flow in lb /sec,

K is the process gain in psi/lb^sec,

T is the process time constant in sec, and

s is the Laplace operator.

Also, K and T are given by

ifi

28.4 A* A0

30

■ ..."

where F0 is the quasi-steady settling chamber pressure in psia,

rh is the quasi-steady nozzle mass flow in lb /sec,

V is the settling chamber volume (15 cu ft),

A* is the throat area (0.105 sq ft), and

T0 is the settling chamber temperature («500oR).

For this process the gain is 2.78 psi/lb /sec and the time constant is

0.225 sec.

b. Pressure Sensor

The pressure sensor consists of three pressure transducers of 0 to

150, 0 to 300, and 0 to 600 psia, switch selectable in accordance with the

range desired. To the control loop the three transducers appear to have the

same characteristic 120 psia/volt; however, the linearity is improved by

switching to a lower range transducer when operating in a lower range. The

transducers are Robinson-Halpern P45 Series, with a rated total error band

of +0.15« of full scale, including hysteresis, linearity and repeatability.

c. Programmer

The programmer is a C6S Model 806. A linear 0 to 10 volt ramp is

generated when a "RUN" command is received. The ramp time is adjustable from

1.3 to 4.3 seconds and is normally set at 3.0 seconds. The control point is

set by adjusting a potentiometer which voltage divides the ramp output;

therefore, regardless of the set-point the output always reaches full scale

in the preselected time.

d. Controller

The first or "outer loop" controller is a CGS Model 671. The

first controller compares the ramp from the programmer to the output of the

31

■ - ■ ■■ ■■---■■'•- ■- .—■-■■

pressure sensor and then acts upon the difference between these signals with

proportional arid reset (integral) action.

The second or "inner loop" controller is a CGS Model 672. The

second controller compares the output from the first controller to the output

of the pressure sensor and then acts upon the difference between these signals

with proportional and reset action. The pressure sensor (P0) signal is input

to the second controller to improve the damping characteristics.

An automatic gain control (AGC) is inserted between the two

controllers to increase the loop gain and compensate for the decreasing

process gain as supply air pressure (Ps) decreases.

e. Servo Amplifiers

The output of the second controller may be manually switched to

either of two CGS Model 661 servo amplifiers, one (SA2) matched to the two

inch control valve and actuator and the other (SA6) matched to the six inch

control valve and actuator. The servo amplifier compares the output of the

second controller to the output of the valve position feedback signal and then

acts upon the difference between these signals to position the valve correctly.

A 60 Hz dither may be added in the servo amplifiers to overcome the static

friction of the valve actuators.

f. Process Control Loop

The complete process control loop is illustrated in Figure 15.

The principal nonlinear element in the loop is the servo valve. The actuator

response is assumed to be a velocity limited capacitance, and the process

response is as previously discussed. The pressure transducer has a linear

response and acts as the feedback element. The "Inner Loop" controller is

32

( ilUHHP ■ -

adjusted to cancel the denominator of the actuator response term, and the

"Outer Loop" controller is adjusted to cancel the denominator of the process

response term.

4. VALVE RESPONSE TESTS

A series of valve response tests were run on each of the control valves

to determine the response time of the actuator and to verify the process time

constant. Figures 16 and 17 show the results of one such test on each of the

valves. The tests were run with the control loop open and by applying a step

command voltage to the input of the servo amplifier, thus causing the valve to

open to a predetermined position in the minimum possible time. The command

voltage was then removed to close the valve in the minimum possible time.

Valve Position (VP) is the output of the valve position indicating unit in

volts and is used only to determine the response time of the control valves.

Ps and Ts are the supply pressure and temperature as measured just upstream

of the control valves. The valve response is well within the specified one

second. The process time constant as determined from each of these tests is

0.23 second, which is in near agreement with the calculated value.

5. WATER HAMMER

From Figures 16 and 17 one can observe the commonly referred to

"water hammer," which is due to momentum exchange of the rapidly accelerating

or decelerating fluid in the pipeline. This can become a serious problem as

mass flows increase in high Reynolds number facilities.

The pressure rise/ ' with the neglect of frictional losses, due to the

rapid closing or opening of a valve and the subsequent change in velocity of

a fluid in a pipe is

33

9

provided the valve closure time is less than the time required for the acoustic

wave to travel the length of the pipeline to the supply reservoir and back.

The time of travel in sec. is

a

where Lp is the length of pipe from valve to reservoir in ft, a is the speed

of sound in air in the pipe in ft/sec, and p is the density of air in the pipe

in lb /ft-'. The frequency of oscillation in Hz in the pipe is

f - 1/t

For a given mass flow the pressure rise in psi becomes

a Arf) _ 1.444 /TST Atfi AP Apg Ap

where

Ap is the pipe cross sectional area in sq'in.,

m is the mass flow in lb/sec, and m

Tj-, is the initial air storage temperature in 0R.

For the Mach 3 facility operating at maximum mass flow of 200 lb /sec, a

supply reservoir temperature of 530OR and a pipe area of 15.52 sq in., the

maximum pressure rise, with no frictional losses, is 428 psi. From Figures 16

and 17 the time for the acoustic wave to travel the length of the pipe and

back is 0.41 sec, which corresponds well with the average length of pipe

between the control valve and the reservoir (250 ft). The distributive nature

of the air storage bottle connections to the pipeline and the number of valves

and elbows between the bottles and the control valve tend rapidly to dampen

the oscillation, and stabilization of the settling chamber pressure occurs in

less than 1.5 sec.

34

■

For design considerations, note that the maximum pressure rise occurs at

the minimum cross sectional area through which the mass must flow; therefore,

cross sectional areas should be designed large to minimize pressure rise. The

pressure rise may be eliminated completely by designing the valve closure

time to exceed the time of travel of the acoustic wave through the pipeline

from the valve to the reservoir and back,

6. CONTROLLER RESPONSE

The effect of the "water hammer" on stabilization of the settling

chamber pressure during a normal controlled run is shown in Figures 18 and 19.

The oscillation affects the pressure ramp; however, it dampens out rapidly,

and stabilization of the settling chamber pressure occurs within the required

four seconds.

The optimum gain and reset rate values were experimentally verified

during initial controller test runs. Satisfactory response is obtained by

maintaining the "outer loop" controller gain at 0.05 volt/volt and reset

rate at 600 repeats per minute. The "inner loop" controller gain must be set

according to the valve size in use, 25 volts/volt for the six-inch valve and

100 volts/volt for the two-inch valve. The "inner loop" reset rate may be

maintained at two repeats per minute for either valve.

7. MODEL SUPPORT SYSTEM CONTROL

The model support system control consists only of the rudimentary

hardware required for a closed loop control system to be added at a later

date. The support sector is moved by a linear hydraulic cylinder which is

connected at a radius of 18.7 inches from the pitch centerline. The cylinder

has a total travel of 6.5 inches, cushions at both ;idr., and will move the

35

rrw/.w.^f-.-rvm:• r*™..™™..'r -T^^^—-.»,—«- -T-.™-.—w^~.wT^™Tr»^7«»:^'i.»r\»-. >,• !*rr*—wrf'"-~>.-^™"r'*-~--fr~?r-<''

sector exactly 20° or +10° from the tunnel centerline. The hydraulic fluid

to the cylinder is controlled by a four-way, center off, solenoid valve which

is in turn controlled by a panel mounted switch. The pitch angle is indicated

by a panel meter which is calibrated in degrees and is driven by the output of

a linear potentiometer operated by the hydraulic cylinder. The error in

indicated angle due to the linear actuation is approximately 0.1%, and to the

panel meter and potentiometer combination approximately +1.5%. The pitch rate

is controlled by a manually adjusted throttling valve in the hydraulic

actuating cylinder return line. It is intended to close the loop between the

feedback potentiometer and the actuator by adding a pitch programmer, a

servo controller, and replacing the solenoid valve and needle valve with a

hydraulic servo valve.

36

lUHJllll , I.MJ I»!.]!!,.!!! .«»..IJIMLU-^-- -

SECTION IV

AERODYNAMIC CALIBRATION OF THE WIND TUNNEL

The aerodynamic calibration of the Mach 3 high Reynolds number facility

whose nozzle and test section dimensions are shown in Figure 20, consisted of

the following series of tests:

1) measurements to determine the lateral and longitudinal Mach number distributions in the test rhombus

2) measurements to determine the magnitude of flow angularity in the test rhombus

3) a limited number of tunnel blockage tests

4) some flow visualization studies to determine the fluid dynamic problems of the facility's start-stop process.

The measurements were made at an average stagnation temperature of 480oR and

three nominal stagnation pressures of 100, 300 and 500 psia. The correspond-

ß fi ß ing free stream unit Reynolds numbers are 18.6 x 10 , 56 x 10 and 93 x 10

per foot, respectively.

1. THE LATERAL MACH NUMBER DISTRIBUTIONS

The lateral Mach number distributions in the test rhombus were

determined from measurements made with a pi tot pressure survey rake, which is

shown in Figure 21. It consisted of a diamond airfoil body with a span of

7.562 inches, a chord of 2 inches, and a maximum thickness of 0.5 inch.

Attached to the leading edge were seven pi tot pressure tubes spaced 1.24 in.

apart. These tubes had an outside diameter of 0.125 in, an inside diameter

of 0.0635 in. and a 30° internally beveled inlet. The outboard pitot tubes

were 1.5 inches long while the centerline tube was a pitot-static probe

2.87 inches long with four orifices located one inch behind the pitot tube

inlet. These four static pressure orifices were 90° apart and were manifolded

37

together to give an integrated average of the local static pressure. All the

rake tubes were made of stainless steel and were silver soldered Inside the

diamond shaped airfoil body. The leads were taken through a hollow sting to

two separate rotary valve and transducer combinations. One of these units

was equipped with a 25 psia variable reluctance transducer and was used to

measure the local static pressure, while the second unit was equipped with a

250 psia variable reluctance transducer and was used to measure the seven

pi tot pressures. The outputs of both of these units were recorded on a

36 channel oscillograph recorder. Other parameters recorded with these units

were the stagnation pressure, stagnation temperature, and the wall static

pressures along the nozzle and diffuser. The pitot rake was positioned in the

facility so that the leading edges of the short pitot tubes were at the

following stations: Xp/xr = 0, +0.345, and +0.690 (as per nomenclature

Indicated in Figure 22). At each of these five longitudinal stations the

survey rake was rolled to four different angles in order to more fully map

the test rhombus Mach number distribution. The roll angle <p was measured from

the horizontal plane in a clockwise direction while looking upstream.

Measurements were made at ^ = 0°, 45°, 90° and 135°, respectively.

Under certain conditions portions of the survey rake extended outside

the test rhombus. An inspection of the geometry Indicates that under the

conditions listed below all the pitot tubes were inside the rhombus.

Vxr <P — Deg

0 0 +0.345 0 +0.690 45

0 45 0 90 0 135

38

■ ■ ■ ■

For the following rake positions the outboard pi tot tube on each side of the

rake was outside the rhombus.

xP/xr ! i * 1 — | ! Deq 1

+0.345 45 i I 90 1 i 135 i

while possibly the two outboard tubes on each side of the rake were outside

the rhombus for the following rake positions.

! XP/xr 1 1 * i — 1 Deg !

+0.690 !

1 ^ 1 45 | 90 \

1 135 1

No data were taken at station x /x = -0.690 with the rake at 4 = 90 since the p r

rake was too large for positioning within the tunnel at this station.

The local Mach numbers were computed from Rayleigh's pi tot equation in

the form

po . r (Y|1)M I^1 r (Y+D y^ P0 " [(Y+l)MJ+2 J [27MMT=T)J (1)

where p is the tunnel stagnation pressure and p' is the measured pitot

pressure. The corresponding Mach number was calculated by computer iteration

of Eq. (1).

Two other methods were used to ,. < the Mach number calculated from

Eq. (1). Since the centerline probe was a pi tot-static probe, the centerline

Mach number was calculated by another method using Rayleigh's equation in the

form

39

■ T

PA = MiMlF L (i^J l"Y-1 (2)

where p' is the measured centerline pitot pressure and p is the locally

measured static pressure. Here too, the Mach number was calculated by computer

iteration of Eq, (2).

The other method for checking the Mach number was carried out through

the use of the energy equation in the following form:

P, 0 - (l ^M2) (3) w

«

where p is the tunnel stagnation pressure and p is the measured nozzle and/

or diffuser wall static pressure. The use of Eq. (3) for such a calculation

implies that the flow is isentropic and that dp/dy = 0 through the tunnel wall

boundary layer.

The lateral Mach number distributions for a nominal unit Reynolds

number of 20.7 x 10 per foot are shown in Figures 23a through 23c. Inside

the test rhombus the measured Mach numbers are within +2% of the design Mach

number of three. In some cases the measured Mach number near the tunnel side

walls was lower than the design value. This decrease in local Mach number is

believed to be caused by very weak nozzle disturbances. These free stream

disturbances are visible in the shadowgraph and schlieren photos shown in

Figures 45 and 46. The corresponding decrease in local Mach number can be

observed in Figure 23a for both fy = 45° and $ = 135°, where it is noted that

the Mach number near the tunnel side walls were calculated to be approximately

2.8, which is about 6.7% below the design Mach number.

40

■ .

..

■ ■

An example where all the pi tot tubes were within the test rhombus is

shown in Figure 23c for x /x = 0. In this case the data indicate that the

Mach number is well within +_2% of the design value. This is true for each of

the four roll angles of <}, = 0°, 45°, 90° and 135°.

The root-mean-square Mach number based only on those pi tot tubes which

were well within the test rhombus for the various planes at x /x =0 are

tabulated below.

4) rms Deq —

0 45 90

135

3.012 2.984 3.023 3.002

The overall root-mean-square Mach number for these specific conditions was

3.005 and was based on a total of approximately 20 measurements.

As was previously mentioned, attempts were made to check these results

by two different methods. The centerline Mach number obtained from Eq. (1)

was checked by calculation of the Mach number with the centerline pitot-static

pressure probe and Eq. (2). The results are shown on most of the curves as a

partially shaded point. In almost every case the Mach number calculated from

Eq. (2) was either equal to or higher than that calculated by the method

dictated by Eq. (1). This difference can be observed in Figures 23a through

23e. Specifically, in Figure 23e for x /x = +0.695 at <£ = 45°, th^ two Mach

numbers were nearly the same; that computed from Eq. (1) was 2.975 while the

Mach number computed from Eq. (2) was 2.990. The maximum difference between

these two Mach numbers can be observed in Figure 23a where x /x = -0.690. p r

In the plane where 4) = 0° it is noted that the Mach number calculated from the

known tunnel stagnation pressure and the measured centerline pi tot pressure is

41

2.975, while that calculated from measurements with a centerline pitot-static

pressure probe is about 3.095, which is about 4% higher than the design Mach

number. It is believed that the difference in the Mach numbers calculated by

these two methods is due to the fact that the static pressure orifices on the

centerline pitot-static probe were not in the optimum location. It should be

mentioned that the points obtained with the centerline probes regardless of

whether Eq. (1) or (2) is used for calculation of the centerline Mach number,

were measured farther upstream by A(x /x ) = 0.118, since the centerline

pitot-static probe was longer than the others and did not agree with the

indicated x /x values shown on Figures 23a through 23e. In Figures 26a, 26b,

and 26c, where the centerline longitudinal Mach distribution is given, the data

have been adjusted by the proper amount.

A second check on the Mach number distribution was made by use of Eq. (3),

these points being indicated as solid points on the curves. In general, the

comparison in the Mach numbers computed by this method and that dictated by

Eq. (1) was excellent. An example of this is shown in Figure 23e, where the

two Mach r.umbers are well within +1% of each other for each value of 0.

Similar results for an intermediate unit Reynolds number are presented

in Figures 24a through 24e for nominal test conditions of p = 300 psia and

T = 480oR, corresponding to an average unit Reynolds number of 59 x 10 per

foot. Here the overall rms Mach number was 2.995. The results for the

highest unit Reynolds number of these tests are shown in Figures 25a through

25e. The nominal test conditions for these data were p = 500 psia at

T = 4910R, with an average unit Reynolds number of 95 x 10 per foot. Here

again the trends and general agreement are as previously described. Under

these conditions the overall rms Mach number was determined to be 3.015.

42

Mvnpvnne ..- _.

The results indicate that the actual rhombus Mach number corresponds very

well with the design Mach number over the unit Reynolds numbers range of the

facility. It also indicates that the lateral Mach number gradients are small.

These conditions are favorable and necessary for good aerodynamic research.

2. THE LONGITUDINAL MACH NUMbER DISTRIBUTIONS

The longitudinal Mach number distributions were obtained from cross-

plotting the lateral Mach number distributions. These data are presented in

Figures 26a through 29c and are divided into four groups depending on the value

of r/As» the normalized radial distance from the tunnel centerline. The values

of r/As are zero (longitudinal centerline), +1, +2, and +3.

In general, all the data points in the test rhombus fall within +2.5% of

the design Mach number. Figures 26a through 26c show the tunnel centerline

longitudinal Mach number distribution for the various free stream unit Reynolds

numbers. In each case there is a slight overexpansion in the upstream portion

of the test rhombus, which is followed by a slight compression through the

downstream part of the rhombus. For example, in Figure 26a, where

Re/£ = 20.69 x 10 per foot, at x /x = -0.8 the mean local Mach number is

about 2.98. As x /x is increased, the local Mach number also increases, p r

reaching a value of 3.025 at x /x = -0.45, indicating „he presence of a weak

expansion wave. As the longitudinal distance x /x is increased further to

x /x = 0.60, the local Mach number decreases monotonically to a value of

2.98, indicating a slight compression in this area of the test rhombus. Once

again it should be noted that these local Mach number changes are very small

and in this particular case they are well witin +1% of the design Mach number.

Similar trends are noted in Figures 26b and 26c, which are the centerline

longitudinal Mach number distributions for Re/£ = 59.2 x 10 and 95 x 10 per

foot, respectively.

43

As we move laterally off the centerline the longitudinal Mach number

distribution varies slightly. For the case of r/As = +1.0, the longitudinal

Mach number distribution is somewhat similar to that for the tunnel centerline

(r/As =0). As can be seen in Figures 27a through 27c, in the upstream portion

of the test rhombus there is a slight overexpansion, ending at a local Mach

number slightly larger than the design value. This region is then followed

by a weak compression, which indicates the rhombus exit Mach number is about

1% lower than the design value.

At r/As = +2.0 and +3.0, the longitudinal Mach number distribution is not

defined as well as in the previous two cases. Figures 28a, 28b, and 28c show

the longitudinal Mach number distribution at r/As = +2.0 for the free stream

unit Reynolds numbers of 20.69 x 106, 59.2 x 106 and 95 x 106 per foot,

respectively. In each case the local Mach numbers vary with longitudinal

distance through the test rhombus in the same manner; i.e., at the upstream

portion of the test rhombus the local Mach number increases monotonically to

a value of approximately 2.983 at x /x = 0, and then becomes essentially

constant from this point to the end of the test rhombus at x /x = +1.0. In

this case the measured mean Mach number is about 0.5% below the design Mach .

number.

Figures 29a through 29c show similar data for the case where r/As = +3.0

for each of the previously mentioned free stream unit Reynolds numbers, As in

the case where r/As = +2.0, there is a slight overexpansion in the upstream

portion of the test rhombus, where the local Mach number increases from 2.994

at x /x = - 0.690 to 3.090 at x /x = -0.340 and then becomes constant at

about 3.01 for the remaining portion of the test rhombus. In the downstream

part of the test rhombus this corresponds to a measured Mach number which is

about 0.35% higher than the design value.

44

3. THE TUNNEL EMPTY MACH NUMBER DISTRIBUTION

The tunnel empty Mach number distribution, based on the assumption that

dp/dy = 0 through the tunnel wall boundary layer, and calculated from the

wall static pressure measurements and the known tunnel stagnation pressures,

is shown in Figure 30. The data were taken at three different unit Reynolds

numbers, namely, 21 x 10 , 59 x 10 , and 95 x 10 per foot. These data are

compared with one-dimensional inviscid isentropic theory in the nozzle only.

The results show that the measured Mach numbers are very nearly independent of

the unit Reynolds number. The data also indicate that the calculated Mach

numbers correspond very well with theory in the subsonic, transonic and super-

sonic portion of the nozzle up to x/x = - 1.15, at which point the Mach number

is approximately 2.65. In the range -1.15 £ x/x <+1.0 the calculated Mach

number is either slightly lower or equal to the design Mach number, depending

upon the unit Reynolds number. Finally, as x/x is increased beyond the test

rhombus, the Mach number shows a continuous decrease with increasing distance.

For example, at the end of the test rhombus the Mach number is 2.94; as we

move downstream from this point to another point midway in the constant area

duct, i.e., x/x = + 4.9, the Mach number decreases to 2.90. In the farthest

downstream portion of the constant area duct (+ 4.9<x/x <+ 9.0) there is a

small difference in the Mach number due to Reynolds number change. T is

difference is considered to be negligible, so that a linear extrapolation to

the end of the constant area duct yields an exit Mach number of 2.82. This

decrease in Mach number from 2.95 at x/x = 0 to 2.82 at x/x = + 9.0 indicates

that a very slight longitudinal pressure gradient exists, which can be attri-

buted to the boundary layer growth ever this distance. This pressure gradient

is considered to be insignificant for purposes of aerodynamic testing in this

45

. . --

facility. The longitudinal pressure gradients for the various stagnation

pressures are as follows:

Po dp/dx

psia psia/in.

108.47 315.34 521.76

+ 1.139 x 10"6

+ 3.305 x 10-f + 5.460 x lO-6

which corresponds to an average Mach number gradient of approximately _3

-1.243 x 10 Mach number per inch.

4. THE CENTERLINE RMS MACH NUMBER

All the Mach number calibration data is summarized in Figure 31 in terms

of the center!ine rms Mach number versus the unit Reynolds number. These

data indicate that the Mach number increases monotonically with increasing

unit Reynolds number. At a unit Reynolds number of approximately 21 x 10 per

foot the average rms Mach is 2.990, while at a unit Reynolds number of

59 x 10 per foot it increases to 2.996. A further increase in unit Reynolds

number to 95 x 10 per foot produces an average rms Mach number of 3.00. The

reasons for presenting this curve are: (1) to summarize the total results of

the calibration in one figure, stressing the fact that all the data are within

+1% of the design Mach number, and (2) to show that the change in the average

rms Mach number is essentially invariant with a change in the unit Reynolds

number.

5. FLOW ANGULARITY MEASUREMENTS

In the design of wind tunnel nozzles, one attempts to achieve a constant

Mach number and uniform parallel flow in the test rhombus. However, this

ideal condition is rarely achieved, since other upstream factors such as the

46

settling chamber design and the nozzle inlet design, can have a significant

effect upon the flow angularity in the test rhombus. Because of this, tests

were conducted as part of the calibration program to determine the presence

of flow angularity in this facility.

The primary models used in these tests were two separate 15 degree

half-angle cones, one mounted on the tunnel centerline and a second one offset

2 inches with respect to the centerline. A wedge model spanning the test

section, having seven pairs of static pressure orifices in the spanwise

direction, was used to extend the range of the test. However, the values of

flow angularity obtained with the wedge model should be considered as qualita-

tive data for reasons which will become obvious later. The pressure level over

the wedge model was made similar to that over the cones by using a smaller

half-angle of 10 degrees. The detailed design features of the wedge are shown

in Figure 32 while those for the cones are shown in Figure 33.

The basic concept of these tests was to determine the flow angularity

from a plot of the difference in upper and lower surface pressure versus angle

of attack. Normally this curve passes through the origin; however, if flow

angularity is present, the angle corresponding to a surface pressure difference

of zero is the angularity present in the flow. Since model orifice irregular-

ities can be a problem, the procedure is to invert the model and repeat the

test. Usually this results in another curve of surface pressure difference

versus angle of attack which has a slope opposite to that for the model up-

right case. Because of orifice irregularities, the angle at which the

pressure difference is zero for the inverted condition may be different from

that for the upright case; therefore, an average flow angularity must be used,

and it is taken as the angle corresponding to the point of intersection of

47

■IVUIfJU^VW^XUftlW ffV'i

these two curves. Examples of some flow angularity curves are shown in

Figures 34a through 34e. These curves are plots of the normalized surface

pressure differences over a cone versus the angle of attack for both the

upright and inverted case. The nominal stagnation pressure was 100 psia at a

nominal stagnation temperature of 470oR; this corresponds to a unit Reynolds

number of 20.7 x 10 per foot. In Figure 34a where x /x = -0.690, when the

model was upright, the flow angularity was + 0.125o(by definition the angularity

is positive when the model is at a negative angle of attack), and for the

inverted case it was + 0.325°. As previously mentioned, this difference in

flow angularity was probably due to orifice irregularities; therefore, an

average value of + 0.225° which corresponds to the point of intersection of

the two curves, was assumed to be correct. This type of accuracy cannot be

obtained in using the multiple pressure orifices on the wedge model, since

inverting the model also transposes the position of each pair of pressure

orifices, except for the centerline pair. As a result, errors due to model and

orifice irregularities cannot be eliminated. The flow angularity data that

follows has been generated in this same manner and will be discussed in

greater detail in the following paragraphs.

The position and geometry of the settling chamber spreader cone was

found to have a signficant influence on flow angularity. In the design of the

facility one of the primary concerns was that due to building constraints on

the piping design the inlet air to the settling chamber would have a high

velocity "spiked" type center core. In order to overcome this undesirable

probability, a perforated spreader cone was installed in the settling chamber

at the upstream end. The purpose of the spreader cone was to provide a full

velocity profile at the downstream end of the settling chamber. As mentioned

48

in the tunnel description, it was designed with a half-angle of 45° and a

porosity of 36%. In order to determine the effect of spreader cone orientation

on the stagnation section velocity profiles, some small scale tests were

conducted. The details of these tests are reported in Ref. 12. In general,

the tests were conducted with a spreader cone having the same half-angle as

the Mach 3 facility; however, the porosity was 40% rather than 36%. The

ratio of air inlet pipe diameter to stagnation section diameter in both cases

was about three, while the stagnation section integrated average velocity in

the smaller unit was 15 ft/sec, as compared to 30 ft/sec in the full scale

facility. There was obviously a Reynolds number or scale effect present

between the two facilities; however, these tests were for the purpose of

obtaining general trends and nothing more. The purpose of these tests was to

determine if the spreader cone would give a more nearly complete velocity

profile in the settling chamber when the apex of the spreader cone was oriented

upstream or downstream. Some of the results are presented in Figure 35; the

three cases investigated are labeled as:

Case A: Turbulence screens without a spreader cone

Case B: The spreader cone apex's pointing upstream plus turbulence screens

Case C: The spreader cone apex's pointing downstream plus turbulence screens.

In each case the velocity downstream of the turbulence screens was measured by

hot wire techniques. As expected. Case A, which was without a spreader cone,

had a high velocity core in the center which extended over the range

-0.45<r/R<+ 0.45, with the centerline velocity equal to 23.0 ft/sec. When

the spreader cone was introduced with the apex upstream (Case B), the constant

velocity region extended from r/R = -0.70 to r/R = + 0.70 and was approximately

49

equal to the centerline velocity, which was 16.9 ft/sec. The spreader cone

was then reversed so that the apex was downstream (Case C), and the tests were

repeated. These data indicated a high peak velocity in the vicinity of

r/R = +0.80, which was greater than Case B, and it had a constant velocity

region extending from -0.60^ r/R <+ 0.60, which was somewhat smaller than

Case B. The centerline velocity for Case C was 13.9 ft/sec, corresponding to

a decrease of 3 ft/sec when compared to Case B. These tests revealed that there

was very little practical difference between cases B and C, but due to the

more extensive region of uniform flow obtained with Case B, it was decided that

the spreader cone in the large facility would be placed with the apex

upstream.

Tests were carried out with both a two-dimensional wedge and a cone to

determine the lateral flow angularity in the test rhombus at x /x = 0. These

tests were conducted at three different unit Reynolds numbers of 20.6 x 10 ,

58.2 x 10 , and 93 x 10 per foot. The data are shown in Figure 36 and

indicate a great change in the flow angularity with increasing lateral

direction.

It varies from approximately -0.4° at z/{b/2) = -1.0, which is the

west wall of the test section, to a maximum value of + 0.80° at the centerline.

It then decreases until it becomes approximately -0.40° at the east wall. It

should be noted that these data were verified by the use of two different

models. This same data have been replotted in Figure 38a with the unit

Reynolds number as a parameter. In general it can be stated that the flow

angularity appears to be essentially independent of the unit Reynolds number

for both the wedge and cone model data. The data shown in Figures 36 and 38a

imply the possible existence in the nozzle of two large counter rotating

50

vortices with upward flow on the tunnel centerline and downward at each side

wall. Reversing the spreader cone so that the apex was downstream yielded the

data shown in Figures 37 and 38b. Thus, the configuration change, which was

permanently adopted, resulted in considerable reduction in flow angularity

distribution to an acceptable limit of approximately + 0.18° and was constant

across the test section. Figure 38b confirmed the previous indication that

flow angularity was essentially independent of the unit Reynolds number.

In order to determine the variation of flow angularity in the longitudinal

direction with the stagnation section spreader cone apex upstream, a shorter

series of tests were conducted with the cone model only. The results are

shown in Figure 39, where the data was taken at M = 2.99 with a stagnation

pressure of 100 psia and a stagnation temperature of 470 R, corresponding to

a unit Reynolds number of 20.6 x 10 per foot. In these tests the parameter

was the longitudinal distance given in its nondimension^lized form as

x /x = 0. x /x = +0.345, and x /x = +0.690. Cross-plots of these data are p r p r - p r - r

shown in Figures 40, 41 and 42. Figure 40 is the longitudinal variation of the

flow angularity at z/(b/2) = + 0.5 (the east side of the test section) while

Figures 41 and 42 are similar plots for z/(b/2) = 0 (the centerline of the

test section) and z/(b/2) = -0.5 (the west side of the test section),

respectively. The results confirm the previous tests and indicate that the

flow angularity is constant at approximately + 0.18° in the longitudinal

direction as well as the lateral direction.

A final flow angularity check was made at the center of the test rhombus

to determine the variation of the flow angularity within the plane containing

the point x /x =0. The test conditions were the same as those given above;

the results which indicate a deviation in the flow angularity as.a function of

51

the polar coordinate positioning of the cone model, are shown in Figure 43.

Tne cone was mounted off-center at a constant distance given by z/(b/2) = + 0.5

and was positioned wichin the plane passing through the point given as

x /x = 0 at various angles denoted by <p. At (J) = 0° the flow angularity was

measured at + 0.24°; it decreases to 0.09° at 0 = 45° and continues to decrease

until it becomes zero at approximately $ = 60°. Then the angularity becomes

negative at cj) = 90° and is -0.09°; from here it increases and becomes zero

at 4) = 135°. As $ is increased from $ = 135° to * = 360°, the flow angularity

increases until it returns to a value of + 0.24° &t <t> = 360°. Superimposed

on Figure 43 is the mean of all the measured centerline values of the flow

angularity, which is + 0.12 , and it checks reasonably well with those

measured in this particular plane. In light of the fact that flow angularity

measurements are difficult to make, the difference in these data is not

considered to be alarming since it is within the expected inaccuracy of the

measuring technique.

6. BLOCKAGE TESTS

Tunnel blockage tests were conducted to determine the criteria on model

size limitations. The models chosen were a sharp-nosed cone and a slightly

blunted cone, both having a half-angle of 15 degrees. In addition, a 10 degree

half-angle wedge was used. The models were placed on the tunnel centerline

with the leading edge at x /x = 0, the rhombus vertical centerline. The

percent blockage was calculated or. th* basis of the model maximum frontal

area plus the exposed part of the support system referred to the physical

cross-sectional area of the test rhombus (A = 65.6 in. ). All tests were

carried out at an angle of attack of zero degrees. The blockage area of both

models was varied by adding or removing bases of different diameters to the

52

jj^W^CW*

i

cone and constant span bases of various heights to the wedge. The cone

percent blockage was changed from 26.11% to 38.19°/, while the wedge blockage

varied from 26.37% to 30.95% of the physical cross-sectional area of the test

section. These values are tabulated at the top of Figure 44, and a plot of

the percent blockage versus Mach number is shown at the bottom of the same

figure. Superimposed are two theoretical curves; one assumes tunnel choking

occurs through a normal shock wave positioned at the model location, and the

second case, which is not too likely, assumes tunnel choking takei place

isentropically. At a Mach number of three, the theory based on tunnel choking

by an increase in entropy indicates the tunnel should start up to a maximum

model blockage of about 28.7% of the physical cross-sectional area of the test

section. In the case of the two-dimensional wedge, tests were performed up

to a blockage of 30.95% with no detrimental effects. For the cone, the

maximum blockage attained without detrimental effect was 33.58%; however, when

the cone blockage was increased to 38.19%, the facility would not start at the

lower stagnation pressures but did go into flow at stagnation pressures

greater than 300 psia. Based on these tests, it is recommended that all models

designed for this facility should not exceed a blockage of approximately 28%

of the physical corss-sectional area of the test section. This should include

the frontal area due to all effects such as sting support, model angle of

attack, flow field probes, etc.

7. REMARKS ON FLOW LIMITATIONS

In order to determine some of the flow limitations encountered during

the operation of the Mach 3 high Reynolds number facility some schlieren and

shadowgraph photographs were taken of the flow field over a wedge. The wedge

had a 10° half-angle and was capable of being pitched to +10 degrees in angle

53

of attack. All the tests were conducted at a stagnation pressure of

approximately 100 psia and a stagnation temperature of 470oR, corresponding to

a free stream unit Reynolds number of 20.7 x 10 per foot. The resulting

schlieren and shadowgraphs, for angles of attack between 0 and + 10 degrees,

are shown in Figures 45 and 46. Because of the high density and large density

gradients present in the flow at these high Reynolds numbers, the shadowgraphs

suffice to show good shock wave details over the model surface, whereas

schlieren will show additional minor disturbances very clearly. In Figures 45a

and 45c, which are schlieren pictures of the wedge at angles of attack of

zero and + 4 decrees, both free stream and minor model disturbances are very

noticeable. In Figures 45b and 45d, the shadowgraphs of the same flow fields,

it is noted that the fred stream disturbances do not appear; however, the

model disturbances are evident but are somewhat reduced in intensity. The

fact that these disturbances are minor can best be observed in an analysis of

the flow field over the wedge as shown in Figures 45a through 45b, where the

angle of attack is zero degrees. If we assume a free stream Mach number of

three, and a wedge half-angle of 10 degrees, calculations using inviscid

theory indicate that the angle of the attached bow shock should be 27.4°, as

compared to a measured value of 27.5 . This corresponds to a local Mach

number ahead of the surface disturbance over the model of 2.5. From this

information and inviscid oblique shock wave theory the angle of the surface

disturbance was calculated to be 23.58°, as compared to a measured angle of

23.4°, indicating that this and similar disturbances are very weak Mach waves.

The primary reason for this test was to determine the starting and

stopping characteristics of this type of facility, sines it is important to

both tunnel operation and proper model design techniques. Prior to these

r)4

- ■

tests, it was assumed that the facility could go into and out of flow by one

of two methods; namely,

1) symmetrically, i.e., with supersonic and/or subsonic flow over both surfaces

2) nonsymmetrically, i.e., with supersonic flow over one surface and subsonic flow over the support system side.

Needless to say, the aerodynamic loads are very different, with the greatest

loads occurring in the latter case. When the model was placed at an angle of

attack of either 0° or + 4° and the tunnel put into operation, a normal flow

field was established over the wedge, as shown in Figure 45. The tunnel was

then shut down and the angle of attack of the wedge was increased to + 10°,

after which an attempt was made at starting the faciltiy. As can be seen in

Figure 46c the flow was unsteady and the normal supersonic flow field was not

established. Specifically, the upper surface of the model was in a supersonic

flow field, which was less than the fully established flow at M = 3, and the

lower surface (which is the model support side) was in a local subsonic flow

field, as witnessed by the presence of a normal shock wave near the model's

leading edge. The tunnel was shut down, the model angl'3 of attack was

decreased to + 7°, and the test was repeated. The results were similar, as

indicated in Figures 46a and 46b. A continuous decrease in angle of attack

during tunnel operation allowed the flow field to become properly established

only after the angle of attack was decreased to +5.5°. Another series of tests

were conducted by first starting the tunnel with the model at zero degrees

angle of attack. Once the flow field was established at these conditions,

the angle of attack was increased slowly while the facility was in operation,

allowing the establishment of the proper flow field at angles of attack as high

as + 10 degrees. The following recommendations are made as a result of the

55

FT-

above tests:

1) For purposes of establishing the flow at all angles of attack,

it is recommended that the facility be started at zero degrees angle of attack,

after which the model may be driven to the desired angle of attack within the

range - 10o<a<+ 10°, The reverse procedure will be detrimental to flow

establishment and the test equipment at the higher angles of attack.

2) All models should be designed to take the loads experienced by

the establishment of a flow field as indicated in Figure 46, i.e., supersonic

flow at M = 3 (even though this is not the case) on the nonsupport side of the

model, and a subsonic flow field on the support system side of the model.

SECTION V

CONCLUSIONS

The overall design objectives for the facility were met and the extensive

testing carried out demonstrated good aerodynamic performance and reliable

mechaniral systems. Some of the more significant characteristics of the

facility are discussed below together with some recommendations to users of

the facilities.

56

1) The lateral and longitudinal Mach number distribution in the

test rhombus is considered to be very good. It varies by approximately +1%

of the design value of 3.0 over the unit Reynolds number range extending from

20 x 106 to 95 x 106 per foot.

2) The lateral and longitudinal flow angularity in the test

rhombus was found to be approximately 0.18° over the entire unit Reynolds

number range of the facility.

3) Blockage tests indicate that models with blockage frontal

areas (including support system) as high as 38% of the physical cross-

sectional area of the test rhombus may be tolerated; however, it is recommended

that 28% should be the upper limit for insurance of proper tunnel operation.

4) Based on some basic start-stop flow visualization studies, the

following recommendations are made:

a. For the purposes of establishing the flow at all angles of

attack the facility should be started with the model at zero degrees angle of

attack, after which the model may be driven to the desired angle of attack.

Starting the tunnel with the model pitched could be detrimental to flow

establishment and to the test equipment, particulary at high angles of attack.

b. All models should be designed to take starting loads based

on a nonsymmetrical flow field, i.e., supersonic flow on the support-free side

of the model and subsonic flow on the support system side of the model.

5) The possibility exists for the facility to be operated at

conditions which could cause condensation of the air in the test section.

Such conditions would normally be encountered only on very cold winter days

when operating at high stagnation pressur ^ for long periods. It is unlikely

that operations would be restricted for more than a few days during the year.

More specific details are given in the Appendix.

57

■

1

6) The noise level of the facility is high with the atmospheric

exhaust configuration and it is essential that established safety procedures

be strictly adhered to. Prior to each series of tunnel runs a positive

verification of the proper functioning of the siren and warning lights should

be made, and during the run personnel should not approach within 50 feet of

the exhaust silencer.

7) The facility vibration level is predictably much higher than

that of a normal supersonic wind tunnel. Care must be taken to periodically

check bolted connections for tightness, and rupture discs should be replaced

at least annually to minimize premature failures due to fatigue,

8) Due to the unheated air supply, the tunnel components are

subjected to temperatures far below the dew point temperature of the air in

the facility building. Subsequent moisture condensation on components follow-

ing a series of runs can produce corrosion problems. Whenever facility

modifications or additions are contemplated, consideration should be given to

the use of corrosion resistant materials.

9) Due to the significant pressure drop between the storage tanks

and the control valves, greater utility of the air supply is achieved by

conducting the highest stagnation pressure tests at the beginning of a series

of runs.

10) For the avoidance of water hammer problems in the air supply

lines, the duration of the stagnation pressure programming ramp should be

maintained at no less than four seconds, and the isolation valve closure

times should never be less than 1 second.

58

11) Closed circuit television should be used to view the test

section windows to insure that no accident occurs as a result of windows

cracking due to thermal effects during a run.

For those who may be involved with the design or operation of a similar

facility, it may be worth noting that the conveyance and control of high

density air at high flow rates and high expansion rates requires a close

examination of conventional design assumptions and practices. In the present

case it was found that the possibilities for water hammer problems, air

condensation problems and low temperature brittle fracture problems were

much greater than expected and other more conventional problems such as noise,

vibration and mechanical stresses were seriously intensified.

59

REFERENCES

1. "AFSC Technical Facility Capability Key," AFSCP 80-3, Air Force Systems Command, 1 July 1965.

2. Schaefer, W. T., Jr., "Characteristics of Major Active Wind Tunnels at the Langley Research Center," NASA TM X-1130, July 1965.

3. "A Compilation of Major Hypersonic Wind Tunnels in the United States," The Boeing Company Document No. D2-5785,

4. "National Wind Tunnel Summary," DOD-NASA, July 1961.

5. Test Facilities Handbook, Vol. 1, Arnold Engineering Development Center, July 1961.

6. "STA Wind Tunnel Listing," Supersonic Tunnel Association, 1966.

7. West, J. E. and Korkegi, R. H., "Interaction in the Corner of Intersecting Wedges at a Mach Number of 3 and High Reynolds Numbers," ARL 71-0241, AD734 540, October 1971.

8. "ASME Boiler and Pressure Vessel Code, Section VIII, Unfired Pressure Vessels," American Society of Mechanical Engineers, 1965.

9. Ferri, A. and Bogdonoff, S. M., "Design and Operation of Intermittent Supersonic Wind Tunnels," AGARDograph No. 1, Advisory Group for Aeronautical Research and Development, May 1954.

10. Azzouz, D. and Pratt, N. H., "Supersonic Turbulent Mixing, Part II, Design and Construction of a Mach 3 Blow-down Wind Tunnel and Its Instrumenta- tion," AASU Rept. No. 282, University of Southampton, May 1968.

11. Pope, A. and Goin, K. L., High-Speed Wind Tunnel Testing, John Wiley and Sons, Inc., 1965.

12. Haugen, R. L.; Moore, D. G. and Wennerstrom, A. J., "Velocity Profile Measurements in a Model Settling Chamber," Internal ARL Paper (available upon request), April 1969.

13. Friedrichs, K. 0., "Theoretical Studies on the Flow Through Nozzles and Related Problems," Applied Mathematics Group AMG-NYU No. 43, Applied Mathematics Panel 82.1R, New York University, April 1944.

14. Nilson, E., "Analytic Correction of a Two-Dimensional Nozzle for Uniform Exhaust Flow," Meteor Rept UAC-15, United Aircraft Corporation Research Dept., January 1948.

15. Baron, J. R., "Analytic Design of a Family of Supersonic Nozzles by the Friedrichs Method," WADC Technical Rept 54-279, AD 55 902, Naval Supersonic Laboratory, MIT, June 1954.

60

16. Burke, A. F., "Turbulent Boundary Layers on Highly Cooled Surfaces at High Mach Numbers," AFASD TR 61-645, ASD Symposium on Aerothermoelasticity, Dayton, Ohio, 1961.

17. Giles, Ronald V., "Fluid Mechanics and Hydraulics," Schaum Publishing Co., 1962.

18. Nagamatsu, H. T.; Sheer, R. E. and Graber, B. C, "Hypersonic Laminar Boundary-Layer Transition on 8-Foot-Long, 10 Cone, M] = 9.1 - 16," AIAA Journal. Vol. 5, No. 7, pp. 1245-1251, July 1967.

19. Pate, S. R., "Measurements and Correlations of Transition Reynolds Numbers on Sharp Slender Cones at High Speeds," AIAA Journal, Vol. 9, No. 6, pp. 1082-1090, June 1971. ~ ^

20. Wegener, P., "On the Experimental Investigation of Hypersonic Flow," Naval Ordnance Laboratory Rept. No. 9629, 1948,

21. Schueler, C. J., "An Investigation of Model Blockage for Wind Tunnels at Mach Numbers 1.5 to 19.5," AEDC TN 59-165, 1960.

61

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63

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66

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79

t— 1 ^

0 I 2 SCALE IN INCHES

A. RAKE SIDE VIEW

B. RAKE PLAN VIEW

Figure 21. Scale Drawing of Mach Number Survey Rakt

80

'

J- 3

4->

o c

c

<c 0) o c s- 3 (/I

CM CM

£ CT

a

si

O - CALCULATED FROM RAKE PITOT PRESSURES

• - CALCULATED FROM WALL STATIC PRESSURES

O - CALCULATED FROM ^ PITOT AND STATIC PRESSURES

-4

-4

-4

"T"

-Gu ■o-

"T"

-G- B.oX-

"T"

-o-

*%/l' 21.61 X I06ft; 2.9 ■

-3 -2

—r

2.8

4> - 0* J.

-T M«

3.01 £1 _a.

•fi

—r

® 2.9 + R,//-22.44 X lO6^'1

! 2.8 2S. X

£L

</> • 45

-3

—r

-2

—r

♦ i

T

Rf/i. X

XIO6^'

-3

—r

-2 -I

-M

3.0

2.9

-2.8

$ • 90*

+ 1

iD-

gy/i- 21.5 6 X 10 6 ft.

4 A~ 3 - 2

WEST WALLl

-I

.a <^ • 135*

+ 1

y/AS

0

X ♦ 2

—r

"O"

♦ 2

♦ 2

—r

"Ö"

2%

X + 3

—r

♦ 4

i 2%

T- i2u

•►3 -»• 4

2%

X •►3

—r -►4

1 2%

ü ♦ 2 +3 t **

I EAST WALL

Figure 23a. Impact Mach Number versus y/As for xp/xr - 0.690 at p0 = 109 psia and T0 = 430oR

82

■

0 - CALCULATED FROM RAKE PITOT PRESSURES

• " CALCULATED FROM WALL STATIC PRESSURES

9 - CALCULATED FROM C PITOT AND STATIC PRESSURES

-4

-4

-4

-4

XT -r ■O

—r IK

«..-•

3

2.9

"T-

■o

R.//" 1993 X 10•ft -L I L -3

—r -2

—r -i

2.8 4> • 0*

±

M

O -GT

O

Rt/i- 21.09 X I06ft' r i i

-3 -2

-r -i

3.0

2.9-fr

•2.8

4-1

T-

^> • 45*

3.0S 0

_Q ^3.

2.9 + R,//- 20.48 X lO6«.'1

JS i ! 2.8 -2

-o- o

S^H

WEST WALL)

-I

♦ I

—r

0

<^ • 90* ±

+ 1

-o-

Rt//- 21.8 8 X 10* ft."' ^^ " '35* 2.8

0

y/AS

+ i

"T"

JL •»•2

—r

0

i ♦ 2

—r

TDr

i. ♦ 2

—r

O

2%

"T-l ±

♦ 3

—r ♦ 4

i_. o_|%

I ♦ 3

—r ♦ 4

"2%

JSL + 3

—r

♦ 4

__t 0 2%

T- 4-2 +3 A 4-4

I EAST WALLl ö

Figure 23b. Impact Mach Number versus y/As for xp/xr - 0.345 at p0 - 111 psia and T0 = 4770R

83



O - CALCULATED FROM C PITOT AND STATIC PRESSURES

-4

-4

-4

-4

-r T

o

M? r" i.oi o-

Rt/i« 21.58 X 10St,'1

3

2.9

-3

—r

-2 2.8

^> • 0* J.

G O- 3.oA.

2.9--

■0-

R9/I' 21.13 X I06ft"1

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</> • 45*

♦ I

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"T"

o «.//• 20.32 X I06ft." -j! I L-

2.9--

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o "cr

-2.8

M

3.0

4> - 90'

+ 1

O 29 , ,

R./i- 18.21 X lO6^ I 9 " '35 . I ' ' 2.8 A-3 - 2 " I C? + I

WEST WALL

V AS

T"

± + 2

—r

-0-

± + 2

—r

■O-

+ 2

—r ■O

■f 2

T

■^■

±

i. 2%

T ■»- 3

—r •► 4

i "GT 2%

± •►3

—r + 4

_LJ — 2%

± •»•3

—r

+ 4

O 2%

T- lU. + 4

EAST WALL

Figure 23c. Impact Mach Number versus y/As for xp/xr 0 at p0 = 111 psia and T0 = 484

0R

84



<1 - CALCULATED FROM £ PITOT AND STATIC PRESSURES

-4

-4

-4

T T

-nr o

,6^-1

3.00

2.9 • Rt/^- 20.78 X 10° ft. j 1 1 2iQ,

-3

—r

-2

—r

-O-

M

3.0

O "S" I

Rt/i- 21.12 X lO6^"1

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2.9 +

•2.8

6 ,.-1 R./i- 20.05X I06ft _li l L_

2.9-

-3

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-M

3.0 -o- "0"

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4> - o*

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^) • 45

-. r-Mr r

■X3 O —Ä 0-

<j> • 90*

+ 1

2.9 R./i" 2 0.34Xl06ft.'i 9-135

■ I I I 2.84 L

A-3 -2 -I 0 +

WEST WALL

v/ AS

T

"Gr

-L + 2

—r

X + 2

—r

•o-

+ 2

—r

+ 2

o i 2 V. T J

± -»• 3

—r ■f 4

i o 2%

+ 3 + 4

T + 3

—r ■♦•4

i 2 V.

O

♦ 3 I + 4

EAST WALL

Figure 23d. Impact Mach Number versus y/As for xp/xr + 0.345 at p0 = 108 psia and TQ = 4770R

85



9 - CALCULATED FROM £ PITOT AND STATIC PRESSURES

-4

-4

-4

-4

T

t-r

H,//' 1905 X 10St

-I My

Q 3oi T"

-i 2.9-

-3

—r -2

—r -i

2.8 4> ' 0*

± + 1

-o- ■o- 0

MI R9/I' 20.36 X lO6^'

I I L_

2.9-•

-3

—r -2 -I

•2.8

o </> • 45*

♦ I

© -O- e hpi 2.9 f

R,//- 20.30 X 106ft.■,

' ' 1 2.8

¥ 4> • 90*

-3

—T

-2

—r -i

M

—O- 3.0

—T

75 ^~2.9

R,//' 20.45 XIO0«."1

-2.8

A o

4> - I35#

t-3 , -2

|WEST WALLj

-I

y/AS

T

J. ♦ 2

—r

■O-

X ♦ 2

—r

ê- j.

♦ 2

—r

■o-

_L. 2% r-

± •*- 3

—r ♦ 4

^L '2%

T-l i

•f3

$

♦ 4

2%

X ♦ 3

—r 4-4

JL. -0- 2%

T- ♦2 +31^4

I EAST WALLl

Figure 23e. Impact Mach Number versus y/As for xp/xr + 0.690 at p0 = 103 psia and T0 = 472

0R

86




-4

-4

-4

-4

-©_

-r

-o-

-1 M^

-O ioi

6^-1 2.9'

R«/i" 60.61 X lO^ft

-3

—r -2

—r -i

2.8

Q 2.9-- R./i- 61.7 3 X I06ft',

& 1 1 2.8 -3

—r -2

—r -i

T

v* X 10 6 ft.*1

-3

—r -2

-M

3.0

2.9

•2.8

M

3.0-- —6

o R./i" 59.48 X I06ft 6 M*-*

2.9--

I - 2 -1—2.8

-I

WEST WALL

4> - o* ±

4> ' A5'

+ 1

4> ■ 135*

0

y/AS

-r O

± + 2

—r

•»•2

T^

♦ 2

^4_- 2%

T x

♦ 3

—r ♦ 4

i 2%

T JX

-»•3 + 4

1 1 I I

4- 2% |

JL + \ <f) > 90*

• i

0 +1 + 2 + 3 + 4

2%

—?T * 4

EAST WALL

Figure 24a. Impact Mach Number versus y/As for xp/xr - 0.690 at p0 = 306 psia and T0 = 468

0R

87




.4

-4

-4

Rt/^- 57.69 X K^ft"' -£- I I -3

O

*.//■ 56-27 XIO6«.'1 Qi ' 1 2.8 -3

—r -2 -I

O 2.9 + Rt//" 59.90 XIO6«."1

■ I I I 2.6

WEST WALL

Figure 24b. Impact Mach Number versus y/As for xp/xr - 0.345 at Po = 316 psia and T0 = 4820R

88

O - CALCULATED FROM RAKE PITOT PRESSURES • - CALCULATED FROM WALL STATIC PRESSURES 9 - CALCULATED FROM £ PITOT AND STATIC PRESSURES

•»-2 +3 f *'4

I EAST WALL

Figure 24c. Impact Mach Number versus y/As for xp/xr 0 at Po = 323 psia and T0 = 4910R

89



O - CALCULATED FROM £ PITOT AND STATIC PRESSURES

-4

-4

-4

-4

—r

-o-

—r

■GT 3.00

6^-1 2.9-

Rt/i- 58.61 X 10° ft. I I I

-3

—r -2

—r

-O- 0 ü

R,/^" 59.38 X lO6^1

-J I L_ -3 -2

•2.8

M

3.0

29:

•2.8

—r

-o-

<j> ■ 0* J.

■f I

—r

-O-

<f> . 45*

♦ I

■O -O- 3.0;

*m/lm 57.14 X ID6»."1 <^ • 90* I I 1 2.8-1 1

-3

—r -2

—T

-I + 1

-M

3.0

—r

O

I O 2.9

*%/lm 57.09 X l06ft.H | 9 ^ 2.8

135

-3 - 2 -I

WEST WALL

-r

JL •»■2

—r

-Q-

X ♦ 2

—r

O

♦ 2

-►2

J. ♦ 3

—r

± + 3

—r

"O"

J. ■f 3

—r

i 2%

T + 4

2%

•f 4

1 2%

+ 4

i 2%

0 ♦!

y/AS

Figure 24d, Impact Mach Number versus y/As for xp/xr = + 0.345 at p0 = 313 psia and T0 = 4830R

+ 3

EAST WALL

90

O - CALCULATED FROM RAKE PITÖT PRESSURES



-4

-4

-4

-4

1 ' ■T ■ I

O vy

1 V^' 59.76 X io« ft;1

L . -3 -2

—r -i

^L

P,//" 59.2 2 X I06ft 6#.-l

-3

—y

-2

—T

•2.8 <j> . 45*

♦ I

J2- .0. G

6 ,.-1

"I 3.0Ä

2.9-• R./i- 61.09 X 10* ft.* t 1 1 2.8

O

<f> • 90*

-3

—T

-2 -I

M

3.0 I ♦ 1

—r

O , 2 9 + Rt/i" 61.47 X 10* ft."*

WEST WALTI

-I -2.8

O

(p " 135*

0

y/AS

+ i

± ♦ 2

—r

-o-

J. >2

—r

i. ♦ 3

—r ♦ 4

-G I— - 2%

T-l X

♦ 3

—r ♦ 4

■&- 2%

♦2 ♦31 +4

lEAST WALL)

Figure 24e. Impact Mach Number versus y/As for xp/xr + 0.690 at p0 = 318 psia and T0 = 4760R

91




-4

-4

. 4

sx -1 M*

6 „.-1 2 9-

R,//- 97.78 X 10* ft. J.

-3

—r -2

—r

2.8

■O"

O 2.9 + R./i-99.43 X lO6/!"'

■2.8 ♦ -^ -2 -1

1 i i

-

yi- —i- X 10 St.'1

1

M

-3

—r -2

3.0--

2.9-■

2.8

"T"

<t> • o#

JL ■fl

-T

■o-

<^ ■ 45#

+ 1

</> • 90*

♦ I

R,/i" 96.27 X I06ft

3 -3 - 2 .2.8

0. - 135*

WEST WALL

0

y/AS

■f i

¥

i

♦ 2

—r

iD-

J. ♦ 2

—r

-^2

-i M m 1 r

^ 3 0i O fS o O-

♦ 2

i3L i 2%

"T J

X ♦ 3

—r ♦ 4

i - 2%

Tt

i-J 2%

J. •»■3

—r f 4

2% r •f 4

EAST WALL

Figure 25a. Impact Mach Number versus y/As for xp/xr - 0,690 at p0 = 518 psia and T0 = 4780R

92




-4

-4

-4

-4

—r -0-

-1 Mr -C —Ö

3.0*

-r

R./^' 88.95 X lC€ff, 2.9-

-3

—r -2

—r -I

-2.8 <f> - o*

X ♦ I

T

G "O"

Ms-

3.0&-

"t/i" 93.00 X lO6^1

-J! i i -3

-T

-2

—r

O—3

2.9+

2.8

"T"

O

-I

if) - 45* 1

T "I 3.0ft-

♦ I

—r

O 3.

2.9 +

0

R,//-90.01 X \06U'i

Jff 1 I 2.8 <f> - 90*

J. -3

—r

-2

"Ü"

-I

-X2L

•n M

3.0 I Q 2.9-- Rt/^" 97.04 XiO6«."1

t-3 , -2

|WEST WAITI

-2.6 4> • 135*

0

y/AS

+1

-T-

•»-2

—r

G

i

♦ 2

—r

"GT

+ 2

—r

0

+ 2

2%

J. ••»• 3

—r ♦ 4

_J_J

L

♦ 3

—r •► 4

_L. 2%

0 -L.

+ 3 + 4

i 0

2%

T + 3 I + 4

EAST WALL

Figure 25b. Impact Mach Number versus y/As for xp/xr - 0.345 at po = 520 psia and T0 = 4910R

93




-4

T

o

R,//" 95.73 X 10* ft

"T" -~r

2.9-

R9/l' 92.78 X I06fr. I I » —I 2.8 ITS Tg^ -I

WEST WALTI

0 4-1

y/AS

-o- i 2%

■f 2 "f3T + 4

|EAST WALL

Figure 25c. Impact Mach Number versus y/As for xp/xr = 0 at Po = 532 psia and T0 = 4990R

94

O - CALCULATED FROM RAKE PITOT PRESSURES • - CALCULATED FROM WALL STATIC PRESSURES 9 - CALCULATED FROM £ PITOT AND STATIC PRESSURES

-4

-4

-4

-4

i ■ ■ -r- n

0 o Ü

Rt/i« 93.94 X ID6 ft.'1

-3

—r -2

—r

R9/l' 95.2 5 X I06ft. -I 29:

-3

—r -2

-2.8

M

3.0

^> • 45*

■XT 2.9

«./,/• 92.62 X 106 ft.'1 \ $ ' SO* Jl 1 1 2.8 J 1

-3

—r -2

—T

-I

3.0

"T"

-O- -^ 0

O 2.9 Rt/i'92.38 X 106ft.H 19 -135

-_j 1 1 2.8-i L

WEST WALL y/AS

± + 2

—r

± -»■2

—r

-o-

•»- 2

T- 4 3 + 4

i o 2'/.

T + 3

—r 4 4

J... — 2%

+ 3 I 4 4

EAST WALL

Figure 25cl. Impact.Mach Number versus y/As for xp/xr + 0.345 at p0 = 517 psia and T0 = 4910R

95


• " CALCULATED FROM WALL STATIC PRESSURES


-4

-4

-4

-4

T —r

-o-

M

6- .3.0j

2.9-• Rt/i- 9 6.19 X ID6 ft."1

J 1 1 2.8,

"T

-o

4> - o* -3

—r -2

—r

o

R,/i- 93.64 X lO6^'1

-J I I

I MA

2.9--

-3

—r -2 -I

•2.8

M

O

</) - 45*

•H

o o 6^-1

Ö-

Rt/i- 97.43 X I06fr.'

3.0

2.9 +

-3

—r -2 -i

-2.8 <^ ■ 90*

+ 1

3. -O- 4

*•/!' &7.0I X 10 ö ft S-.H 2.9-

E - 2 •2.8

O

$ ■ 135*

WEST WALL

0

y/AS

T

_L + 2 —r

-O-

± + 2

—r

-o-

+ 2

—r

^x.

+ 2

T 0_L

2%

+ 3

—r

± ♦ 3

—r 0

± + 3

—r

-»- 4

i 2%

T + 4

1 2%

T 4 4

^_ 2%

T + 3 I + 4

EAST WALL

Figure 25e. Impact Mach Number versus y/As for xp/xr + 0.690 at p0 = 523 psia and T0 = 4860R

96

T i

/ /

/ /

o / /

/ /

/

/

3.06

3.04 ■-

a:

2 v

<I

z

¥ 00 ■ - ^^r-

2 98--

Ui 2 96

^ 2.94--

2.92 --

i. 2.90-

OESIGN MACH NUMBER

THR\?"[ -FAIRED THR"Ü» DATA

SYM. i ^ — I DEC 0 0 9 45 0 9 0 A 13 5

-L 1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 06 0.8

*

Xp/XR

o

Figure 26a. Center!ine Mach Number versus Longitudinal Distance for po = 100 psia, To = 4680R, where Re/£ = 20.69 x 106 per foot

97

. i T i %

o o

3.06

3.04-

T - /

- /

- /

i u 1 DEG. O 0 9 45 0 9 0 ^ 135

X -L JL 1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0

Xp/XR

Figure 26b. Centerline Mach Number versus Longitudinal Distance for po 300 psia, T0 = 480.06oR, where Re/£ = 59.2 x 106 per foot

98

T i ^

1 a

3.06

304-

tt ui

3N Q

I i u <

UJ

roo

2.98 -

/ a: w 2.96 z UJ o • 2.94

2.92 -

2.90-

0ESI6N MACH NUMBER

^ O

FAIRED THRU DATA 2 SYM. i <t> — IDE

if G. 0

AS. 90

13 5

± JL 1

■1.0 -0.8 -06 0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0

Xp/Xp

Figure 26c. Centerline Mach Number versus Longitudinal Distance for p0 500 psia, T0 = 489.05oR where Re/£ = 95.0 x 106 per foot

99

T I ^

1 0

306

3 04-

3.02 -

/ fi sa:

DESIGN MACH NUMBER

■0—00 3.00

3

x 298 o <

. 2 96 +

CD

2.94--

2.92 -

J. ± 2.90-

Ä

^ 0 ■FAI ÄD THRU DATA

o ^^

"-1 SYM. * o

DEG. 0 0

5 45 A ? 90 A 13 5 A

-L J. 0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/XR

Figure 27a. Mach Number versus Longitudinal Distance for r/As = + 1.0 at Po = 100 psia, To = 468.40R and Re/£ = 20.69 x 106 per foot

100

3.06

T I t

i DESIGN MACH NUMBER

O

O

r /0

/ I

I <s>

'S

! 298-- ^

< Z , 2 96-

2.94--

2.92 -

I ^ L ± 2.90-

0

■FAIRED THRU DATA

IS SYM. || </> i DEC O 0 9 45 ? 90 AA 135

A

J. 1 1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/XR

1.0

Figure 27b. Mach Number versus Longitudinal Distance for r/As = ± 1.0 at Po = 300 psia,T0 = 480.06oR and Re/£ = 59.2 x 106 per foot

101

T0 3.06

T i

3.04 ■ ■

3.02 - -

Nt

•DESIGN MACH NUMBER

^ Wx 0 3.00

O/ 3

2 98 - -

/ /

/ /

x u <

• 2 96 4-

2.94 - -

2.92 - -

J. ■L2.90-

¥ S

0 O

•FAIRED THRU DATA

^^^ o

SYM. * DE6. O 0 e 45 ÖA 90

| A | 135 1

± ± 1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/Xp

1.0

Figure 27c. Mach number versus longitudinal distance for r/As = ± 1.0 at Po = 500 psia, T0 = 489.05oR and Re/£ = 95.0 x 106 per foot

102

T I %

1°

3.06

3.04-

OESIGN MACM NUMBER

/ /

/ 6 I

h / / /

< i G i ■^2.90-

^

2 94-

2.92 -

f-

SYM. 1 * 1 DEG. o t"~o~i ( ~Ö| TJ 45 o 1 90 ! A 1 i 135 1

± ± ± 0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/XR

Figure 28a. Mach Number versus Longitudinal Distance for r/As = + 2.0 at p0 = 100 psia, T0 = 468.40R and Re/£ = 20.69 x 106 per foot

103

T 1

3.06

3.04 -

DESIGN MACH NUMBER

/ a /

/ O

<

Q , 2 96 I O 5

FAIRED THRU DATA

t I

I I

h / /

/

i L

2.94--

2.92 --

± J. 2.90-

SYM. 1 </> — 1 DEG. 0 0 9 45 0 9 0 A 13 5

o i. J.

.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/XR

Figure 28b. Mach Number versus Longitudinal Distance for r/As = ± 2.0 at p0 = 300 psia, T0 = 480.06oR and Re/£ = 59.2 x 106 per foot

104

T i •/

i ©

306

3.04-

3 02 -

O

Ul m 3.00

I 1

DESIGN MACH NUMBER

o

-o- ^"

o z __4> T—o Ö- o 0 ^-FAIRED THRU DAT^

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DATA Ä

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SYM. * DEC

O 0 9 45 § 90 A 135

J 2± ± ± •10 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0

Xp/Xp

Figure 28c. Mach Number versus Longitudinal Distance for r/As = ± 2.0 at Po = 500 psia, T0 = 489.05oR and Re/Z = 95.0 x 106 per foot

105

i T 1

306

3.04 - -

. 2 96

JL 2.90- Ö

O

DESIGN MACH NUMBER

FAIRED THRU DATA

i SYM. * 294 A-

I O DEG.

0 D 45 O D 90

2.92- Ä 1 135

J. ±

Figure 29a. Mach Number versus Longitudinal Distance for r/As = + 3 0 at Po = 100 psia, TQ = 468.40R and Re/Z = 20.69 x 106 per foot

10 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0

Xp/XR

106

306

T i

/ /

/

_ / /

/

_/ /

/

/

i o <

, 2.96--

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2.92

-L2.90-

FAIRED THRU DATA

SYM. f DEC [

1 O 1 0 9 ! 4 5 5 ) 9 0

! A I 113 5

10 -08 -06 -04 -0.2 0 0.2 04 06 08 10

Xp/XR

Figure 29b. Mach Number versus Longitudinal Distance for r/As = ± 3.0 at p0 = 300 psia, T0 = 480.06oR and ReA = 59.2 x 106 per foot

107

3.06

T i ^

1

s / 0

J. •2.90

0

DESIGN MACH NUMBER

FAIRED THRU DATA

SYM. | 4> 1 DEG. O 0 5? 45 6© 90 A 135

1 -L 1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8

Xp/XR

Figure 29c. Mach Number versus Longitudinal Distance for r/As = ± 3.0 at Po = 500 psia, TQ = 489.05° and Re/£ = 95.0 x 106 per foot

108

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O.IZ'X 45 CHAMFER

OTOT

A. WEDGE PLAN VIEW

1.00" 0.75"

B. WEDGE SIDE VIEW

Figure 32. Scale Drawing of Flow Angularity Wedge

111

UJ

< tit It

tf) Q. 3

O o

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UJ z o u

c o

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Figure 34a. Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = - 0.690 for p0 = 103.6 psia with Settling rhamber Spreader Cone Tip Downstr&am

113

Figure 34b. Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = - 0.345 for p0 = 100.6 psia with Settling Chamber Spreader Cone Tip Downstream

114

-1.6

A-UPRIGHT 0-INVERTED

L_

Figure 34c. Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = 0 for p0 = 100.6 psia with Settling Chamber Spreader Cone Tip Downstream

115

Q IN DEGREES

A-UPRIGHT O-INVERTED

Figure 34d. Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = 0.345 for p0 = 95.0 psia with Settling Chamber Spreader Cone Tip Downstream

116

Figure 34e. Normalized Surface Pressure Difference versus Angle of Attack at xp/xr = 0.690 for p0 = 95.4 psia with Stagnation Section Spreader Cone Tip Downstream

117

SPREADER CONE

± T

CASE A

TURBULENCE SCREENS

i --U

CASE B CASE

d2-*U

r/R

1.0

0.9

0.8

07

0.6 -

0.5 -

0.4 -

0.3

0.2

0.1

0

i—i—i—i—i—i—r

j i L

Figure 35. Velocity Profile in a Model Settling Chamber with and without a Spreader Cone

118

i 1 1 r- 0-2* WEDGE (0/2-10') Q-CONE (0/2« 15')

-0.4 - -

■0.8

M • 3.01 ffi ..-I Rt/i- 93.42 X IGT ft.

_J I I L

J -0.8

Figure 36. Flow Angularity versus Lateral Distance at xp/xr = 0 (Rhombus Center) with the Spreader Cone Tip Upstream

119

T

O - 2d WEDGE (d/2 ■ 10*) Q-CONE (9/2 - 15*)

0.8 8

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J 1 ., 0.8

CAST WALL

+-oH h 02 0.4 0.6 04 1.0

04 -L M - 2.99 Rî' 20.62 X 10* ft"*

± ± J L

Figure 37. Flow Angularity versus Lateral Distance at xp/xr = 0 (Rhombus Center) with the Spreader Cone Tip Downstream

120

SYMl M 20.62 58.18 93.42 20.62 58.18 93.42

WEDGE

-I 1 ' L-0.6

A. SPREADER CONE TIP UPSTREAM

1 1 0.6

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t7

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Z/(b/2) -*-

_L J L

B. SPREADER CONE TIP DOWNSTREAM

Figure 38. Unit Reynolds Number Effect on the Flow Angularity at xp/xr = 0 (Rhombus Center)

121

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J -0.4

0.2

"Cr

4 0 Z/lb/2

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r E^T WALL A

6 0.8 1.0

Figure 39. Flow Angularity versus Lateral Distance with Longitudinal Distance as a Parameter for pn = 100 psia. To = 470oR, M = 2.99 and Re/£ = 20.6 x 10^ per foot

122

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BLOCKAGE TEST RESULTS

| MODEL d h AAt/A |

INCHES INCHES

j CONE 4.28 .2611 j

4.53 .2866 |

4.75 .3035 [

5.00 .3358 |

* 5.41 .3819 |

|2fl WEDGE 1.492 .2637 |

1 1.745 .2866 |

i 2.000 .3095 | L ,." ,_ „

XD/XR - 0 AND A- 65.6 IN.'

0.8

Ui o < * u o

♦- 2 It! U or UJ Q.

0.6

0.4 -

0.2 -

NORMAL SHOCK WAVE CASE

ISENTROPIC CASE

M - MACH NUMBER

Figure 44. Tunnel Blockage versus Mach Number

127

(a) SCHLIEREN AT Q = 0 (b) SHADOWGRAPH AT (3*0

(c) SCHLIEREN AT Q -+ 4 (d) SHADOWGRAPH AT Q =+4°

Figure 45. Schlieren and Shadowgraphs of Flow Over a Two-Dimensional Wedge at M = 3.0 and Re/£ = 20.66 \ 106 per foot for a = 0° and a = 4°

128

(a; SCHLIEREN AT a = +7 b) SHADOWGRAPH AT Of = +

(c) SCHLIEREN AT Q '+10

figure 46. Schlieren and Shadowgraphs of Flow Over a Two-Dimensional Wedge at M = 3.0 and Re/' = 20.66 x 10fi per foot for a - 7° and a + 10l

29

.

APPENDIX

BASIC DESIGN CONSIDERATIONS

The starting point for the design of any wind tunnel is a knowledge of

the aerodynamic performance requirements which arise from the particular testing

need. These requirements are usually expressed in terms of the desired test

section flow conditions, model sizes, and run times. Figure A-l is a design

flow chart which illustrates how this basic input information is used, and

modified when necessary, to arrive at firm design parameters which form the

basis of the detailed design of the facility. The chart is not complete in

the sense that there are always special conditions and constraints which

significantly affect the design process. In the present case these included

the size and location of the building available to house the facility, and the

very common constraint of limited funds. This appendix deals with the basic

design process, as distinct from the detailed design process, and discusses

some of the key points.

1. INPUT DESIGN PARAMETERS

a. Test Section Mach Number

A test section Mach number of M^ = 3.0 was required with no

provisions for changes in Mach number, other than by replacement of the nozzle

blocks at some future time to accommodate nozzles up to Mach numbers of 4.5,

There will always be some variation in test section Mach number due to changes

in the nozzle wall boundary layer over the facility operating range but for

basic design purposes such variations are not usually significant. At the high

Reynolds numbers under consideration here this is particularly true, since the

boundary layer displacement thickness is very small, and changes in it are

proportionately small. The Mach nuuâr was therefore held constant at M^ = 3.0

throughout the design process.

131 Preceding page blank

b. Desired Model Sizes

Models on the order of one foot long were envisioned as being

sufficient to accommodate the necessary instrumentation and to permit external

probing with reasonably sized probes without producing severe probe disturbance

effects. Probable model geometries included right-circular cones, wedges,

flat plates, delta wings and corner flow models.

c. Desired Reynolds Number Range

The Reynolds number range of the tunnel was to be as large as

possible, and preferably sufficient to give both laminar and turbulent flow

conditions over the models. Based upon Mach 3 transition data available at /1 o\

the time,v ' it appeared that transition took place at Reynolds numbers from

1.5 x 106 to 2.0 x 106 for planar flow fields. Later work by Pate^ ^ showed

that corresponding transition for conical flow fields at Mach 3 would occur

at Reynolds numbers higher by a factor of 2.2 to 2.5. Figure A-2 shows how

these data translate into freestream unit Reynolds number requirements for

transition to occur at a certain point on a model. For example, in order to

have transition on a cone completed no further than 0.5 inch from the apex,

a freestream unit Reynolds number of at least 1.2 x 10 per foot would be

required. It should be noted that Figure A-2 can only serve as a rough guide,

such as tunnel size, tunnel freestream turbulence, model surface roughness,

and tunnel wall boundary layer noise.

d. Desired Run Times

Run times of at least 60 seconds were sought to allow the possibility

of detailed flow field probing of the models, and to permit the use of a

pitch-pause model support system. Due to the relatively short response times

132

associated with high Reynolds number testing, it was anticipated that many

runs would average only from 20 to 30 seconds.

2. CALCULATED PARAMETERS

a. Tunnel Stagnation Conditions

(1) Stagnation Temperature, T0

For high Reynolds number testing, where high freestream

densities are sought, the test gas is usually unheated unless there is a danger

of the gas liquefying upon expansion through the nozzle. At high Mach numbers

the accommodation of this aspect often dominates the facility design, but at

lower Mach numbers it is not a major problem. This is not to say that it can

be ignored, even at a test section Mach number of 3, especially if the tunnel

is to operate during winter from a high pressure gas storage system exposed to

the elements. This was the case here, and since one of the ground rules for the

construction of the facility was that the air supply be unheated, changes in

stagnation temperature during and between test runs must be accepted, together

with the resulting changes in Reynolds number. The temperature changes are

brought about not only by changes in the ambient air temperature but just as

significantly by the cooling which takes place in the storage tanks as the air

expands polytropically during the depressurization process (a function of mass

flow, run time and tank volume), and by the cooling (Joule-Thomson effect)

associated with the throttling through the pressure control valve. Since the

ARL high pressure air storage tanks do not contain heat storage materials,

long run times may cause appreciable temperature changes as a result of the

expansion of the air in the tanks, and will be most serious at the higher

Reynolds numbers, where the mass flows are largest. Joule-Thomson effects

will be more pronounced at the lower Reynolds numbers, due to the greater

133

pressure drop across the pressure control valve. These cooling effects will be

partially offset by the transfer of heat from the walls of the storage tanks

and piping containing the air, and by the frictional heating associated with

the pressure drop through the supply lines to the tunnel. In summary then,

factors which influence the temperature of the air delivered to the nozzle in-

clude the ambient temperature, the initial storage pressure, the rate of

depressurization of the air in the storage tanks, the stagnation pressure

required for the run, the run time, and the position of the run within a

sequence of runs. While such temperature changes cannot be controlled, and are

difficult to estimate with any reasonable precision, they can be easily

monitored by continuously recording the tunnel stagnation temperature.

Figure A-3 shows the liquefaction limits for air based upon Wegener's equation^ '

for static conditions, and assumes an isentropic expansion to Mach 3 from the

indicated total conditions. The parameter AT represents possible degrees of

cooling below the initial air storage temperature (TST) caused by expansion and

throttling effects. The solid line curve corresponds to the case where the

tunnel stagnation temperature (T0) is equal to the initial air storage

temperature, i.e., AT = 0. The broken line curves show how the static

liquefaction limits change with increasing amounts of cooling. The broken line

curves are in fact displaced to the right of the solid curve by an amount on

the abscissa corresponding to the indicated value of AT. Since preliminary

estimates showed that values of AT up to 100oF were entirely possible, it is

clear that the effect of an unheated air supply has to receive consideration

before the operating pressure range can be completed. Figure A-4 shows to

what extent the unit Reynolds number depends upon the tunnel stagnation

temperature, and gives the percentage change in unit Reynolds number per degree

Rankine change in temperature for the stagnation temperature range of interest.

134

(2) Stagnation Pressure, P0

With a fixed Mach number nozzle and an unheated air supply,

the freestream Reynolds number is primarily governed by changes in the tunnel

stagnation pressure. From the previous consideration of desired model sizes

and boundary layer conditions, freestream unit Reynolds numbers between

approximately 1.5 x 10 and 1.2 x 10 per foot appeared to be necessary, with

a probable stagnation temperature range of 400oR to 500oR. Figure A-5 gives

the variation of freestream unit Reynolds number with stagnation pressure for

these two stagnation temperatures. At a stagnation temperature of 500oR,

stagnation pressures between 8 psia and 685 psia are necessary to produce the

required unit Reynolds number range, whereas at a stagnation temperature of

400oR stagnation pressures between 6 psia and 485 psia are sufficient.

However, to avoid condensation of air in the nozzle, operation at stagnation

temperatures of 400oR would have to be limited to stagnation pressures of less

than 400 psia (see Figure A-3).

b. Tunnel Pressure Ratio, A

The probable maximum and minimum pressure ratios required to start

and run a wind tunnel at various Mach numbers are shown in Figure A-6, which

is based upon data given in Reference 11. For a test section Mach number of

3 it appears that a pressure ratio no greater than 5.40 is necessary to start

the tunnel, and a pressure ratio as low ac 2.75 may be sufficient to keep the

tunnel in flow. Based upon achieving a essure recovery equivalent to that

obtained behind a normal shock at the test section Mach number, a value of

} = 3.05 would be required. In order to exhaust to standard atmospheric

pressure, a stagnation pressure of 79.4 psia would be necessary to start the

flow, and a stagnation pressure of 40.4 psia would be necessary to maintain

135

the flow once established. This shows that the lower end of the full Reynolds

number range previously identified cannot be achieved without a vacuum exhaust

or auxiliary pumping capability.

c. Test Section Size

The test section size is primarily governed by the sizes and types

of models envisioned for testing which have been previously identified. With

the discussion restricted to closed test sections with square cross sections,

the most important model parameters are length and thickness (or diameter),

which can be conveniently coupled into a parameter called the fineness ratio

which is defined as the ratio of model length to maximum model thickness or

diameter; i.e., L/tm or L/dm. For a particular test section size, the limita-

tions on the model size are those due to the model bow shocks being reflected

from the tunnel walls back onto the model, and those due to the inability to

start or maintain the flow in the test section because of excessive blockage

or restriction of the flow caused by the model and the model support system.

In general, models having large fineness ratios are limited by bow shock

reflection considerations, whereas models having small fineness ratios are

more likely to be limited by blockage considerations. Figure A-7 applies to

bow shock reflection limitations and shows how the ratio of maximum model

length to test section half-height, L /hjs, varies with the fineness ratio

for some common model configurations. The figure can be used to find either

Lm or hys when the othar is specified. The limiting case of a flat plate of

infinitesimal thickness is indicated for reference purposes. Figure A-8

applies to blockage limitations and shows how the ratio of maximum model

diameter to inviscid test section area, d /A-, varies with test section Mach

number for blunt models. The theoretical curve is based upon an analysis

136

which equates the unblocked area with that necessary to "swallow" a normal

shock during the tunnel starting process. In practice, duv co the losses in

total head associated with the system of model shocks, the actual flow area

required is larger than the theoretical value which assumes an isentropic flow

downstream of a normal shock at the test section Mach number. The experimental

curve shown in Figure A-8 was taken from Reference 21, and indicates that for

testing blunt bodies at Mach 3 the allowable blocked area must be reduced from

the theoretical value of 28% of the inviscid test section area to 12X, at

least during starting. Once the flow is established, blunt body blockage

values on the order of 28% can probably be tolerated. For slender models with

weaker shock systems higher blockage values should be possible (as evidenced

by the calibration tests reported in Section IV). Figure A-9 presents maximum

model sizes as functions of test section size for bow shock reflection limited

cases, and Figure A-10 presents the same information for blockage limited cases.

Vertical lines corresponding to test section sizes of 8 inches x 8 inches and

12 inches x 12 inches are indicated for reference purposes.

d. Mass Flow

The mass flow required to operate a facility at a particular Mach

number is proportional to the test section density ano area. It is an

important parameter since it describes the demands made on the high pressure

air supply system and the exhaust system, considerations which usually limit

the size of a tunnel. Figure A-ll gives the mass flow rate per unit test

section area as a function of the tunnel stagnation conditions. For the tunnel

sizes of interest, the maximum mass flow rates are on the order of hundreds of

pounds per second. For example, an inviscid test section of 8 inches x

8 inches operated at Mach 3, with a stagnation pressure of 500 psia and a

137

Stagnation temperature of 500oR, would require a mass flow of 180 lb/sec. A

12-inch x 12-inch test section operated under the same conditions would require

405 lb/sec. Providing and controlling such mass flows is not a simple matter,

especially for run times on the order of 60 seconds, and the subject is

discussed further in connection with the air supply system and the exhaust

system.

e. Exhaust System

The purpose of the exhaust system is to accept the air discharging

from the wind tunnel diffuser, and to safely discharge it to the atmosphere.

This must be done without producing an excessive back pressure on the

diffuser, which would cause the tunnel flow to break down due to an insuffi-

cient pressure ratio. Exhaust systems are usually of the atmospheric or

vacuum types. An atmospheric system employs no pumping devices, and therefore

limits the lowest Reynolds numbers attainable with a facility. The primary

concern with an atmospheric exhaust system is to diffuse the air to a reason-

ably low velocity (on the order of 100 ft/sec) and to provide silencing. A

vacuum exhaust system usually employs <: number of staged vacuum pumps separated

from the tunnel by a large vacuum tank, although this can be dispensed with

if the pumps are able to accept the entire tunnel mass flow. Ejector systems

are sometimes employed in place of mechanical pumps, but for the high mass

flows of interest here they are not an economical proposition. Requirements

for atmospheric and vacuum exhausts are discussed in more detail below.

(1) Atmospheric Exhaust

For reasons of economy and operational simplicity an

atmopsheric exhaust system is desirable. Obviously this becomes more difficult

to achieve with increasing test section Mach number but at Mach 3 it is a

138

definite consideration. Referring to Figure A-6, one can see that the probable

minimum tunnel pressure ratio required to maintain flow is 2.75, which gives a

minimum tunnel stagnation pressure of 40.4 psia at standard atmospheric

pressure. With a stagnation temperature of 500oR this would produce a free-

stream unit Reynolds number of 7.08 x 10 per foot which is higher than the

minimum Reynolds number desired for laminar flow testing.

In addition to the question of Reynolds number range, another

significant factor associated with a blowdown tunnel which has an atmospheric

exhaust is the noise problem. Noise originates both inside the system, due to

internal turbulence created by such devices as constrictions, control valves,

corners, and shock waves, and outside the system due to the turbulent mixing

occurring in the shear layer between the exhaust jet and bounding atmosphere.

The former source of noise will cause the containing pipes and ducts to

vibrate, producing noise radiation from the entire tunnel, while the latter

will radiate from the discharge opening and points downstream. Once the

internal flow conditions have been made as aerodynamically smooth as possible,

a normal endeavor in wind tunnel design, the only way to handle internal noise

is to minimize atmospheric transmission by accoustical treatment of the internal

or external surfaces. However, this is not normally done, since the primary

source of noise is more likely to be from the exhaust discharge. The intensity

of the jet noise is largest at the discharge opening because the shear velo-

city is highest and gives rise to the most violent eddies which also have the

highest frequencies. As the eddies are convected downstream their kinetic

energy is converted to potential (pressure) energy, and they propagate as

quadrupole noise sources. The eddies formed at the shear layer downstream of

the discharge opening become progressively milder, larger, and of lower

139

frequency. To minimize the jet noise it is clear that the first consideraiion

must be to diffuse the flow as much as possible prior to discharge. In this

way the energy available to the noise producing mechanism is reduced. This

has the obvious advantage of minimizing the required tunnel pressure ratio,

but at the same time it increases the size and expense of ducting and components

The addition of a silencer is the second consideration, and presents the

necessary compromises between effectiveness and pressure drop and between size

and less expense. Another factor which cannot be overlooked is the effect of

weather conditions on th.. design, operation, and life of the system. The

result of the silencer considerations is usually determined by pressure drop

and cost limitations. This points toward the use of silencers of the absorp-

tion type, which employ straight-through acoustically transparent perforated

annular tubes, or rectangular sandwich panels, containing noncorrosive material

having a high sound absorption coefficient. Figure A-12 illustrates the size

of the silencer required for conditions of interest with a silencer inlet

velocity of 100 ft/sec. Since the density at the silencer inlet is essentially

constant regardless of flow conditions, the size of the silencer required is

directly proportional to the tunnel mass flow.

(2) Vacuum Exhaust

The only time a vacuum exhaust system is normally considered

is when an atmospheric exhaust system would unacceptably restrict the operating

range of the tunnel or when it is important to reduce tunnel starting loads.

For the present case it was seen above that the minimum unit Reynolds number

attainable with an atmospheric exhaust was 7.08 x 10 per foot, for the stated

conditions, whereas numbers on the order of 1.5 x 10 per foot were previously

discussed as being desirable for completely laminar conditions on a flat

140

plate model 12 inches long. Figure A-13 shows thp vacuum volume required to

operate at a Reynolds number low enough to achieve this condition for 60 seconds,

assuming the vessel to be completely evacuated prior to the test run. The

calculations were made assuming that the pressure in the vessel at the end of

60 seconds would be equal to that necessary to cause flow breakdown at the

tunnel pressure ratios indicated. For operation at the highest Reynolds

numbers, a 60 second run would pressurize the vessel above atmospheric

pressure if an automatic venting system were not incorporated. Clearly the

size and expense of a vacuum exhaust system are directly affected by the tunnel

mass flow.

The jet noise problem associated with a vacuum exhaust is less

than that for an atmospheric exhaust, since the exhaust jet mixing process and

shock system are confined to the vacuum vessel, although the vessel itself will

transmit noise to the atmoshpere. Trie internally produced noise originating in

the wind tunnel proper will not differ from that of the atmospheric exhaust

case, except for effects caused by differences in exhaust ducting configurations.

f. Air Supply

The air supply system, which consists of air storage tanks and

associated distribution piping, must be able to deliver the required quantity

of air at the proper pressure and temperature for the required length o.-' time.

The system characteristics which determine whether this is possible include

the air storage volume and related pressure, and the size, length and geometry

of the distribution piping. The demands placed upon the system are expressed

in terms of the tunnel mass flow rate, stagnation pressure, and run time. The

tunnel mass flow is a key parameter, since it affects the minimum storage

pressure required at the start of a run, the pressure drop through the piping.

141

the depressurization rate during the run, and therefore the temperature of the

air delivered to the tunnel. In addition to the piping pressure losses, the

pressure drop across the tunnel pressure control valve and tunnel flow con-

ditioning equipment affects the minimum storage pressure required at the start

of a run. It is convenient to express these losses in terms of the tunnel

stagnation pressure, since this largely determines the mass flow rate for a

cold flow tunnel of given size. Control valves typically have pressure drops

ranging from 10 to 100% of the controlled pressure, so that for estimating

purposes a pressure drop of 0.75 p0 would be a realistic allowance, and would

probably be sufficient to accomodate the tunnel flow conditioning loss as well.

Losses through the air supply piping can be roughly estimated at 0.75 p0.

More nearly exact estimates of piping losses are discussed later, but at this

point it is seen that there is a requirement for the storage pressure at the

end of a tunnel run to be at least 2.5 p0. The decrease in storage pressure

due to the cooling caused by the expansion of the air in the storage tank will

aggravate the problem but the conservatism of the estimates should accommodate

this aspect. Figure A-14 shows the storage volume required for three 60 second

runs, at the indicated pressures, as a function of tunnel size. The initial

storage pressure has been taken as 2200 psia, rather than the maximum system

pressure of 3000 psia, since the air supply and recovery systems at ARL serve

several major high pressure users. The final storage pressure has been taken

as two and one half times the tunnel stagnation pressure for reasons discussed

above. The strong effect of tunnel stagnation pressure on air storage require-

ments is readily apparent from the figure.

3. OUTPUT DESIGN PARAMETERS

The foregoing calculations were made to see how difficult it would be to

meet the requirements expressed by the input design parameters. Performing

142

these calculations determines the possible range of design parameters, but the

final selection of output design parameters still depends on value judgments

and personal philosophies. The output design parameters listed below resulted

from numerous discussions between interested parties, and reflect a compromise

between technical ideals concerning tunnel size and performance and practical

considerations of cost and operability. Supporting arguments for the final

choices are given where appropriate.

a. Test Section Mach Number

As indicated previously, the test section Mach number was held

constant at M^ = 3.0 throughout the design process.

b. Test Section Size

It is quite simple to show that for mass flow conservation the

maximum Reynolds number based on model length iz obtained with the smallest

practical model size and the largest practical stagnation pressure. Since the

model size is proportional to the test section size, mass flow consideration

dictate the use of the smallest practical tunnel. Mass flow has a direct

bearing on operating costs, and upon the initial •investment required when an

adequate air supply does not already exist. Apart from air supply cost

considerations, the cost of the tunnel itself is roughly proportional to the

test section area, so that there are strong incentives to build a small tunnel.

Quite early in the preliminary design process it was established that the

smallest acceptable test section size would be 8 x 8 inches, based upon model

instrumentation requirements. However, since the attainment of a certain

model Reynolds number depends on the available pressure, as well as tunnel

size, it is clear that test section size is a parameter which should be varied

in the study. In the case of an infinitely large air storage volume, it is

143

obvious that the maximum model Reynolds number would rise in direct proportion

to tunnel size, since the maximum supply pressure would be independent of

consumption. However, in the case of a finite storage volume there is an

optimum size tunnel which will yield maximum model Reynolds numbers.

Figure A-15 presents the results of calculating maximum attainable model Rey-

nolds numbers, for a range of test section sizes, based on the maximum stagnation

pressures possible with an air supply volume of 8,250 ft (Figure A-14).

It i1; interesting to note that the model Reynolds number is an optimum for a

10 x 10 inch tunnel, although it is only 3.6' higher than that obtainable with

an 8 x 8 inch tunnel. Since the mass flow for the 10 x 10 inch tunnel would

be 29.6 higher and the test section area would be 56 larger, both significant

cost factors, it was decided to design the tunnel on the basis of an 8 x 8

inch inviscid test section size. For comparison purposes Tables A.I and A.II

show the testing capabilities of an 8 x 8 inch tunnel and a 10 x 10 inch

tunnel. The tabulated information is based upon Figures A-l, A-9, A-10, and

A-15. Although the information relates only to zero incidence testing, it

does illustrate that an 8 x 8 inch tunnel should be quite capable of testing

a wide variety of models under well developed turbulent flow conditions, and

to a lesser extent under laminar flow conditions.

c. Stagnation Temperature

Since the tunnel was not to include a heater for the supply air, the

stagnation temperature depends upon the ambient temperature and process

conditions. Ambient temperature at the facility site (Wright-Patterson AFB,

Ohio) ranges from a minimum -200F to a maximum of 100oF. Process conditions

could cause the stagnation temperature of the air to be as much as 100oF below

ambient, so that the stagnation temperature range could extend from 340oR to

144

5600R. However, practical considerations of air condensation effects, and low

temperature limitations associated with ASME coded pressure vessels, require

ttat operation at the lower temperature be subject to certain restrictions,

further details on the ASME code limitations will be given later.

d. Stagnation Pressure

The maximum stagnation pressure for design purposes was chosen to

be 570 psia. This corresponds to a design point unit Reynolds number of one

hundred million (10 ) per foot at a stagnation temperature of 500 R, and is

consistent with the maximum pressure deduced from air storage volume considera-

tions summarized in Figure A-14. The minimum stagnation pressure possible is

basically a function of tunnel pressure ratio requirements and exhaust condi-

tions. For atmospheric exhaust conditions Figure A-6 suggests a minimum

running pressure of approximately 40 psia, although pressures on the order of

80 psia may be required for flow initiation. For vacuum operation, with an

available volume of 100,000 ft^. Figure A-13 illustrates that there should be

no problem in achieving Reynolds numbers as low as one million (10°) per foot.

The limiting factor for low pressure operation with a vacuum exhaust can be

expected to be the stability of the pressure control system.

e. Run Time

The original design goal of maximum run times on the order of

60 seconds is quite feasible. Factors which will influence the actual run

times employed will include the time required for flow establishment, the time

required for temperature stabilization (if important), instrument response

times, model temperature stability (if important), flow visualization require-

ments (if appropriate), and air consumption.

145

f. Exhaust System

The decision was made to construct the facility with a vacuum

exhaust capability while allowing for atmospheric blowdown operation as well.

While a wide Reynolds number range was desired, the principle motivation for

having a vacuum exhaust capability was exhaust noise attenuation. Due to the

unusually high flow densities and flow rates, there was considerable concern

about the effectiveness of low pressure drop silencers. As an interim measure

it was therefore agreed to exhaust the tunnel into the 100,000 ft vacuum

sphere, and to reconsider this decision after gaining some operational experi-

ence of noise and vibration problems. As discussed in Section II of this

report this did become necessary.

g. Performance Envelope

The output parameters discussed above are conveniently summarized

graphically in Figure A-16. The various zones indicated on the figure are

defined as follows:

Zone A - Operation with atmospheric exhaust and tunnel

stagnation temperature within normal ambient range. No special restrictions.

Zone B - Operation with vacuum exhaust and tunnel stagnation

temperature within normal ambient range. No special restrictions.

Zone C - Operation with atmospheric exhaust and tunnel

stagnation temperature below normal ambient range due to air supply polytropic

expansion and throttling effects. Special consideration necessary to insure

that tunnel wall temperature does not fall below 440CR (-20oF).

Zone C - That portion of Zone C in which air condensation

effects are possible. Operation in this zone not desired.

146

Zone D - Operation with vacuum exhaust and tunnel stagnation

temperature below normal ambient range due to air supply polytropic expansion

and throttling effects. Special consideration necessary to insure that

tunnel wall temperature does not fall below 440oR (-20oF).

147

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INPUT DATA BAoED ON j NATURE OF TESTING |

MCD' K« Lin • L^m\n '

'min

INPUT DATA BASED ON

NATURE OF FACILITY

P.., ■ Pot • PST. .VST . V^ | APCV, Lp , Ap ,TST |mjn j

CALCULATE REQUIRED UNIT i FREE STREAM REYNOLDS |

NUMBER RANGE : i

R,eoirn = f (R« |m" •Lm)

1

CALCULATE REQ'D STAGNATION j

PRESSURE RANGE |

| mot / | ma« 1 'olmln='V"ta, lmin> !

To |3 , Moo)

ASSUME T0 = TST TO START

A. REDUCE R.m„

AND/OR INCREASE tmmin

NO / Pt S PSTI \ ■• N. 0mo«

1 [YES

© ®

Figure A-l. Design Flow Chart

150

REDUCE R. ■mo

AND/OR

INCREASE L„

AND/OR

REDUCE t„,„

®

CALCULATE REQ D INVISCID

TEST SECTION AREA

I m01' , / \

CALCULATE REQUIRED

NOZZLE THROAT AREA

A* |2" -■ f (Mo,, A, r^

CALCULATE REO'D NOZZLE

MASS FLOW RATE

m ma. , ] mai mo» 1 ma« \

LZ CALCULATE/ESTIMATE DROP

IN TOTAL PRESSURE FROM

STORAGE TANKS TO TUNNEL

SETTLING CHAMBER

AP s f (fn t.

APcv,Lp,Ap)

CALCULATE STTRAGE PRESSURE

AT END OF TUNNEL RUN

^ P I ♦ AP > n o I mm I mm,/

Figure A-l, Continued

151

©

CALCULATE STAGNATION TEMP. IN SETTLING

CHAMBER imoa |ma>

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ADJUST INPUT VALUES

AT I man

Imin

ESTIMATE DROP IN TOTAL TEMP FROM STORAGE TANKS TO SETTLING

CHAMBER I mo« / I man

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LIMIT VALID STAGNATION

TEMPERATURE RANGE

BY INCREASING T0 min

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CALCULATE FREE STREAM

STATIC TEMPERATURE

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CALCULATE FREE STREAM

STATIC TEMPERATURE REQ'D

TO AVOID AIR CONDENSATION

'«cond Imin

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YES

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Figure A-l. Continued

152

®

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NO (VACUUM EXHAUST)

CALCULATE EXHAUST PRESSURE

AT END OF TUNNEL RUN

PUT P. >2 Imm = P'"

REDUCE R,m

AND/OR

INCREASE L„

AND/OR

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YES

CALCULATE AVAILABLE

TUNNEL PRESSURE RATIO

OUTPUT DESIGN DATA

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DETAILED DESIGN OF WIND

TUNNEL COMPONENTS , CONTROLS,

AND INSTRUMENTATION

Figure A-l. Concluded

153

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STREAMWISE MODEL DISTANCE (INS)

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154

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2.AT-INITIAL AIR STORAGE TEMP-T0

300 400 500 600

INITIAL AIR STORAGE TEMPERATURE (0R)

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155

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156

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157

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TEST SECTION MACH NUMBER, M®

Figure A-6. Tunnel Pressure Ratio Requirements

158

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159

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THEORY,SECOND THROAT ANALYSIS

EXPERIMENT(REF.2I)

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160

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161

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162

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163

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INVISCID TEST SECTION AREA {\WC)

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164

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INVISCID TEST SECTION AREA.A(IN^)

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165

15

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100 200

INVISCID TEST SECTION AREA.ACIN^)

Figure A-14. Variation of Air Storage Volume Requirements with Test Section Size and Operating Pressure

166

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'oo O "CVJ in 2

_J L_™~__J 1

200 400 INVISCID TEST SECTION AREA(INZ)

Figure A-15. Maximum Model Reynolds Number Attainable as a Function of Inviscid Test Area

167

O O

LJ Ü-

cn 10 LJ GD

CO Q _l O

>- Lü

H

3

< LÜ cr

UJ UJ a:

8

10'-

10 -

Moo = 3.0

570

ATMOSPHERIC EXHAUST

CONDITIONAL OPERATION

PSIA

DESIGN POINT

ATMOSPHERIC EXHAUST

VACUUM EXHAUST

CONDITIONAL OPERATION

B

VACUUM EXHAUST

5 PSIA

300 400 500 600

TUNNEL STAGNATION TEMPERATURE (0R)

Figure A-16. Predicted Tunnel Performance Envelope

168

LIST OF SYMBOLS

a speed of sound of air

A cross-sectional area of test section

A* cross-sectional area of nozzle throat

Ai inviscid test section area

Ap cross-sectional area of pipe

b semi-span of test section

Cv control valve flow coefficient

d base diameter of blockage cone

d] inlet diameter of model settling chamber

02 outlet diameter of model settling chamber

dm maximum model diameter

d£ dy

pressure gradient normal to nozzle axis

-j^- longitudinal pressure gradient

f frequency of oscillation

g gravitational constant

Gj process transfer function

h base height of blockage wedge

hjj height of test section

K process gain

K] process gain of controller

K2 process gain of servo-controller

K/\ process gain of actuator

L length of model

Lmax maximum length of model

169

LIST OF SYMBOLS (CONT'D)

Ln length of nozzle measured from throat

Lp length of a section of pipe

if) mass flow rate

m quasi-steady nozzle mass flow rate

M local Mach number in flow field

Mrms root-mean-square Mach number based upon local Mach number measurements

Mw Mach number at tunnel wall

Mx local Mach number at distance x

pc static pressure in test section

Pd stat c pressure in diffuser

pe static pressure in exhaust duct or vacuum sphere

p0 quasi-steady settling chamber pressure

P local static pressure

P] supply line static pressure at inlet to isolation valve

Pz supply line static pressure at inlet to isolation valve

Pcond static pressure at onset of air condensation

P0 stagnation pressure or settling chamber pressure

PQ measured pitot pressure

Ps supply line static pressure at inlet to control valves

P5J pressure in storage tanks

Pw measured nozzle or diffuser wall static pressure

r/R normalized radial distance in model settling chamber

r/As normalized radial distance from nozzle axis to impact probe

ReLn Reynolds number based on nozzle length

170


Rem Reynolds number based on model length

Rex local Reynolds number at distance x

Re/£, Reoo 'free stream unit Reynolds number

Rn radius of circular arc section of nozzle contour

s Laplace operator

t time period for one oscillation

tm maximum model thickness

Tc static (wall) temperature in test section

Tcond static air temperature at onset of air condensation

T0 stagnation temperature or settling chamber temperature

Ts static temperdture of air at inlet to control valves

TST temperature of air in storage tanks

u mass-averaged velocity in model settling chamber

U local velocity in model settling chamber

V volume of sett1 irj chamber

VP valve position

VSJ volume of air storage tanks

X, x axial distance from nozzle throat

Xp, xp axial distance of pi tot tube head measured upstream from nozzle exit station

XR, xr axial distance of upstream extremity of test rhombus measured from nozzle exit station

y nozzle ordinate distance from nozzle axis

y/As normalized ordinate distance from nozzle axis to impact probe

z spanwise distance from nozzle axis

171


AA^ total blockage area

Ap pressure differential or pressure drop

As pi tot probe spacing on Mach number survey rake

Aj temperature differential due to throttling and expansion = l^j - T0

AV change in fluid velocity in pipe

A angle of attack

c, flow angularity

6* local displacement thickness of boundary layer at distance x

f displacement thickness at nozzle exit

Y ratio of specific heats for air

> tunnel pressure ratio

ij Mach angle

$ roll angle of survey rake measured clockwise from horizontal plane looking upstream

p density of air

half-angle of wedge or cone model

T process time constant

r] process time constant of controller

'2 procesr time constant of servo-controller

[ß process time constant of actuator

0 included angle of flow angularity cone or wedge

o«. shock angle

172

U.S.Grn/ernmem Printing Offi.c: 1975 6b7-020/483

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AD-A009 138 DESIGN AND CALIBRATION OF THE ARL MACH …