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~A7OAOBA511 KAMAN AVIOSNE BURLINGTON MA F/6 21/5F URTHER EVALUATI'ON OF BLAST TESTS OF AN ENGINE INLET. CU)

MiAR 79 J R RUETENIK. R F SMILEY. M A TOMAYKO DNAO-78-C-0239

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DNA 4994F

Oo FURTHER EVALUATION OF BLAST

o TESTS OF AN ENGINE INLET

:Kaman AviDyne

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IBurlington, Massachusetts 01803

31 March 1979

Final Report for Period 1 April 1978-31 March 1979

CONTRACT No. DNA 001-78-C-0239

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-F1RTHER VALUATION OF BLAST TESTS OF AN" FinalRpar-for PeriodENGINE ADLET, ADDR 1 Apr 78-31 MarOCT-------------.... . .. . . A-E RFQBA

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19. KEY WORDS (Continue on reverse side if necessary and Identify by block number)

Aircraft Engine ShockBlast Experimental Test Shock TubeB-i Inlet SubsonicComputer Studies Pressure Wind Tunnel

20. ABSTRACT (Continue on reverse side If necessary and Identify by block number)

A test program was conducted to simulate blast wave intercepts with a scaledaircraft engine in subsonic flight. Initial results were reported in DNA4590F. This report presents additional test data and evaluated results.

Jumps in total pressure at the engine face produced by blast interaction withthe blastward and leeward inlets are evaluated with respect to effects offour test variables: inlet weight flow, Mach number, shock overpressure and

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0. ABSTRACT (Continued)

intercept angle. The total pressure rose more rapidly in the blast inlet thanthe leeward inlet, following blast intercept, and by a larger amount. In manytests the rise in the leeward inlet was followed by a rapid fall-off, indi-cating possible flow separation in the inlet walls.

The two-dimensional Blast Induced Distortion (BID) code provides good repre-sentation of effects essentially two-dimensional and inviscid, such as ramp,cowl and engine-face pressures of the blastward inlet.

Blast wave reflection from the engine is expected to be a potential cause ofdistortion through boundary-layer retardation. The effect on engine-facetotal pressure is examined with respect to the four test variables. Improvedcalculations of the effect of the engine-reflected blast wave on the boundarylayers in the inlets, made using the generalized NASA-Lewis BLAYER code,indicate that boundary-layer separation would result.

UNCLASSIFIEDSECURITY CLASSIFICATION OF THIS PAGE(17ten Date Entered)

Conversion factors for U.S. customaryto metric (SI) units of measurement.

To Convert From To Molt irls Rv

angstrom me'ter. 1n 0,t 1 00 X S I n1

atmosphere (normal) kln pas..ii (kl'a) 1.011 215 XC EC +2

ha kiln pa.,.1 (kPs i.0,10 '0 X EC 4.

barn mcer Wn) 1.000 000 C EC -."

British thermal unit (thermochemieal) Joule (J) 1.0Sf. 350 XC E +3

calorie (thermochemical bole, (J1) 4.184. 00n

cal (thermochemical)/cm2 met., joule/m, (MI/.') 4.18. 00n X E -2

curie gtig bcqocrel )Cq) .Ion 000 X r +i

degree 0-ngi0 radian (cad) 1.745 329 IC I -2

degree Fahrenheit degree kelvin (K) Wt f + 45q.67l/l.S

electron volt joule (J) 1.602 19 XC F -19

erg jool e (J) 1.000 000 X F. -7

ergt/.econnd watt (W) 1.0 no 0 OM x F -7

(not meter In) 1.0~ 00 " fton x -1

loot-pound-force joule (.3) i.lS5 ale

gallon (U.S. liqu.id) mterd We) 3.705, 412 X F -3

inch meter (.n) 2.%',n 000 X EC -2

jerk joule (J) Loon0 000o x E 49

JOUle/klOgR. (J/kg) (radiat ion dose

,absorbed) Cray ICV),* 1.000 000

kilotons teraJ.ules 4.1R)

hip (1000 lbf) ne..on WN 4..0 222 X E 43

hip/inch2 (kni) kilo ponil (hI~a) 6.894 757 XC E +3

ktap ncwton-occond/m2

(N-sis') 1.00n 000 x F +2

micron metr,* (oI.001 000 C IC -6

mil Pltr In)'. 2 n 00X F -$

mile (international) meter In I.1.09 34. X IC +3

ounce kilogram (hgl 2.814 952 X IC -2

pound-forte (lhf avoirdupoin) newton (N) 4.416A .2

pound-force inch newton-meter (Nm) 1.129 SO9 X E -1

pound-force/inch newton/meter (N/in) 1.7i1 26A X IC +2

pound-forccffont2 kiln pascal WI.) 4.788 026 X IC -2

pound-force/inch7

(psi) kilo Pascal (kPa) 6.894 757

pound-mann (Ibis avoirdupois) hilogra. (kig) 4.535 924 X IC -1

pound-mass-root2 (moment of Inertial kilogr3m-meter2

(khgm',) 4.2114 il X IC -2pound-mans/oot

1 ktilram/meter

(kg/n1) 1.601 A!, f, x F +1

red (radiat ion dose absorbed) Cray (rCy)* 1.000 000 C IC -?

roentgen rolooh/kilgream (C/k;) ?.579 760 XC F -4

shake s"eond (s) 1.000 000 IC IC -8

plus kilogram (kgt) 1.459 390 C IC +1

tort (wi Hg, 0* C) kilo Pa,ral NkP.) 1.311 22 X IC -1

*The becquorel (8q) in the Si unit of radinsettvitl; I Bq - I event/s.**The Cray (Cy) In the SI unit of absorbed radiation.

A mnts Complete listing of conversions may he found in "Metric Practice Guide E 380-74,American Society for Testi-ug and Materials.

PREFACE

This work was performed by the AviDyne Division of the Kaman

Sciences Corporation for the Defense Nuclear Agency under Contract

DNA-O0l-78-C-0239. CAPT J. Michael Rafferty of the DNA Shock Physics

Directorate served as technical monitor.

Dr. J. Ray Ruetenik of Kaman AviDyne was the project leader under

Dr. Norman P. Hobbs, Technical Director of KA. Mr. Robert F. Smiley

performed engineering functions.

Appreciation is expressed to CAPT Rafferty for his continuing

interest and support of this program.

ACCESSION for

NTIS White SectionOoc Buff Secit

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DISTIO)UII(AYROWIJ CMEDis? I - 4ii. /or 3PECI.

2

Lei= N

TABLE OF CONTENTS

Section Page

1 INTRODUCTION --------------------- 15

2 RESPONSE OF MEAN TOTAL PRESSURE AT ENGINE FACE - - - - 20

2-1 GENERAL FEATURES OF EFFECT ON MEAN TOTALPRESSURE --------------------- 20

2-2 MACH 0, 0 LB/S FULL-SCALE WEIGHT-FLOW ------ 232-3 MACH 0.55, 235 LB/S FULL-SCALE WEIGHT FLOW- - - 232-4 MACH 0.55, 350 LB/S FULL-SCALE WEIGHT FLOW- - - 242-5 MACH 0.70, 300 LB/S FULL-SCALE WEIGHT FLOW- - - 252-6 MACH 0.70, 350 LB/S FULL-SCALE WEIGHT FLOW- - - 262-7 MACH 0.70, 350 LB/S, ±50 YAW ----------- 282-8 MACH 0.85, 300 LB/S FULL-SCALE WEIGHT FLOW- - - 282-9 MACH 0.85, 350 LB/S FULL-SCALE WEIGHT FLOW- - - - 29

2-10 MACH 0.90, 350 LB/S FULL-SCALE WEIGHT FLOW- - - - 29

3 GENERAL FEATURES OF BLAST AND OPERATIONAL EFFECTSON ENGINE-FACE MEAN TOTAL-PRESSURE SIGNATURE ----- 56

3-1 EFFECT OF BLAST SHOCK OVERPRESSURE -------- 563-2 EFFECT OF BLAST INTERCEPT ANGLE --------- 58

3-3 EFFECT OF WEIGHT FLOW -------------- 583-4 EFFECT OF MACH NUMBER -------------- 59

4 SPECIFIC EFFECTS OF BLAST AND OPERATIONAL VARIABLES ON

ENGINE-FACE MEAN TOTAL PRESSURE ------------ 116

4-1 VARIATION IN FIRST PEAK OF MEAN TOTAL PRESSUREFOR BLASTWARD INLET --------------- 1164-1.1 Effect of Incident Total-Pressure

Increment --------------- 1164-1.2 Effect of Test Variables on Engine-Face

First-Peak Ratio, ------------ 1174-1.2.1 Shock Overpressure Effect- - - 1174-1.2.2 Shock Intercept Angle Effect - 1174-1.2.3 Weight Flow Effect ------ 1184-1.2.4 Mach Number Effect ------ 118

4-2 VARIATION IN SECOND PEAK OF MEAN TOTAL PRESSUREAT ENGINE FACE FOR BLASTWARD INLET- ------- 1184-2.1 Effect of Shock Overpressure ------ 1194-2.2 Effect of Intercept Angle ------- 1194-2.3 Effect of Weight Flow --------- 1204-2.4 Effect of Mach Number --------- 120

4-3 SUMMARY FOR BLASTWARD INLET ----------- 120

3

TABLE OF CONTENTS (CONCLUDED)

Section Page

4-4 FEATURES OF MEAN TOTAL PRESSURE FOR LEEWARDINLET -- --------------------- 121

4-5 FLOW SEPARATION WITHIN LEEWARD INLET -- ------ 1214-6 CONCLUSIONS- ------------------- 125

5 COMPARISON OF THEORY AND EXPERIMENT -- --------- 139

5-1 BLAST INPUT CONDITIONS -- ------------- 1395-2 INLET PRESSURES- ----------------- 1415-3 ENGINE FACE PRESSURES- -------------- 141

5-3.1 Blastward Pressures- ---------- 1425-3.2 Lpeward Pressures- ----------- 142

5-4 DISTORTION AT ENGINE FACE- ------------ 1435-4.1 IDC- ------------------ 1435-4.2 IDR- ------------------ 1455-4.3 IDL-------- ----- - - --- -- -- -- - - 1455-4.4 IDA- ------------------ 1455-4.5 IDT- ------------------ 146

5-5 CORRELATION SUMMARY- --------------- 146

6 EVALUATION OF LATE TIMlE LARGE DISTORTION VALUES- - -240

7 REFLECTED SHOCK-BOUNDARY LAYER INTERACTION- -- ---- 242

7-1 SHOCK-BOUNDARY LAYER CALCULATION -- -------- 243

8 CONCLUSIONS- -- -------------------- 250

REFERENCES -- --------------------- 252

APPENDIX -TRANSDUCER CALIBRATION EVALUATION- -- --- 253

4

LIST OF ILLUSTRATIONS

Figure Page

1.1 Sketch showing typical shock wave pattern and

transducer locations in the inlets ---------- 18

1.2 Engine face transducer locations ----------- 19

2.1 Typical engine-face mean total pressures. Run 8.

Mach 0.70, W2R=300 lb/s (nom), Aps=5.0 psi, =98 deg.- 30

2.2 Engine-face mean total pressures. Run 1. Mach 0,W2R=0 lb/s (nom), APs=2.7 psi (nom), 0=79 deg. 31

2.3 Engine-face mean total pressures. Run 2. Mach 0.55,

W2R=235 lb/s (nom), APs=4.7 psi (nom), f=91 deg. - - - 32

2.4 Engine-face mean total pressures. Mach 0.55,

W2R=350 lb/s (nom) ------------------ 33


W2R=300 lb/s (nom) ------------------ 35

2.6 Engine-face mean total pressures. Mach 0.70,W2R=350 1b/s (nom) ------------------ 38


W2R=350 lb/s (nom) ----------------- - 44

2.8 Engine-face mean total pressures. Mach 0.85,W2R=300 lb/s (nom) ------------------ 46

2.9 Engine-face mean total pressures. Mach 0.85,W2R-350 lb/s (nom) ------------------ 51

2.10 Engine-face mean total pressures. Mach 0.90,W2R-350 lb/s (nom) ------------------ 54

3.1 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.70.

Weight flow 350 lb/s (nom). Shock Tube 1 ------- 60

3.2 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 1 ------- 61


5

LIST OF ILLUSTRATIONS (CONTINUED)

Figure Page





3.8 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Shock Tube 1 ---------- 67


3.10 Effect of blast shock overpressure on engine-face mean

total pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Shock Tube 2 ---------- 69





3.15 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.55.Weight flow 350 lb/s (nom). Blast shock overpressure

3.8 psi (nom) --------------------- 74

6


Figure Page


2,7 psi (nom) --------------------- 75

3.17 Effect of blast intercept angle on engine-face mean

total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock overpressure

2.9 psi (nom) --------------------- 76


4.9 psi (nom) --------------------- 77



4.4 psi (nom) --------------------- 78

3.20 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.85Weight flow 300 lb/s (nom). Blast shock overpressure

4.9 psi (nom) --------------------- 79


4.0 psi (nom) --------------------- 80



3.7 psi (nom) --------------------- 81


total pressure in leeward inlet at Mach 0.70. Weightflow 300 lb/s (nom). Blast shock overpressure

2.7 psi (nom) --------------------- 82


total pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Blast shock overpressure

2.9 psi (nom) ------ ------ ---------- 83

3.25 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.70. Weight

flow 350 lb/s (nom). Blast shock overpressure

4.9 psi (nom) --------------------- 84

7


Figure Page

3.26 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 300 lb/s (nom). Blast shock overpressure4.9 psi (nom)- -------------------- 85



3.29 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.90. Weightflow 350 lb/s (nom). Blast shock overpressure3.7 psi (nom) --------------------- -88

3.30 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 1. Shock overpressure 2.8 psi (nom) ------- 89

3.31 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. ShockTube 1. Shock overpressure 3.3 psi (nom) -0------

3.32 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 2. Shock overpressure 2.7 psi (nom)------ - 91


3.34 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. ShockTube 2. Shock overpressure 3.9 psi (nom)- ------ 93



8


Figure Page

3.37 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. Shock

Tube 3. Shock overpressure 4.4 psi (nom) ------- 96

3.38 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 1.

Shock overpressure 2.8 psi (nom) ----------- 97






Shock overpressure 5.0 psi (nom)- - ---------100


Shock overpressure 3.9 psi (nom) -i---------- 101




Shock overpressure 4.3 psi (nom) ---------- - 103



3.46 Effect of Mach number on engine-face mean total

pressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure

2.8 psi (nom) -------- ------------- 105

3.47 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure

3.7 psi (nom) ---------------------- 106

9


Figure Page

3.48 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom) - --------------- 107

3.49 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom) --------------------- 108

3.50 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom) --------------------- 109

3.51 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shock overpressure4.2 psi (nom)- - ----------------- - 110

3.52 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure

2.8 psi (nom)- -i- - - ---------------- 111

3.53 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom) --------------------- 112

3.54 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom)- ------------------- 113



4.1 Sketch illustrating first and second peaks of engine-face mean total pressure for blastward inlet ----- 127

10


Figure Page

4.2 Increment in mean total pressure at blastward engine-face versus incident total pressure increment ----- 128

4.3 Engine-face peak ratio versus incident overpressure- 129

4.4 Engine-face peak ratio versus blast intercept angle- - 131

4.5 Engine-face peak ratio versus weight flow- ------ 133

4.6 Engine-face peak ratio versus Mach number ------- 136

4.7 Sketch illustrating features of engine-face mean total

pressure for leeward inlet -------------- 138

5.1 Blast input characteristics for Run 39 (Part 544)- - - 147



5.4 Comparison of theoretical and experimental time

histories of ramp pressures for Run 39 (Part 544)- - - 150


histories of cowl pressures for Run 39 (Part 544)- - - 151

5.6 Comparison of theoretical and experimental timehistories of ramp pressures for Run 40 (Part 619)- - - 152


histories of cowl pressures for Run 40 (Part 619)- - - 153

5.8 Comparison of theoretical and experimental timehistories of ramp pressures for Run 18 (Part 624)- - - 154

5.9 Comparison of theoretical and experimental timehistories of cowl pressures for Run 18 (Part 624)- - - 155

5.10 Comparison of theoretical and experimental timehistories of engine face total pressures for Run 39

(Part 544), blastward (outboard) inlet -------- 156

5.11 Comparison of theoretical and experimental timehistories of engine face total pressures for Run 39(Part 544), leeward (inboard) inlet ---------- 164


Figure Page


histories of engine face total pressures for Run 40(Part 619), blastward (outboard) inlet -------- 172


histories of engine face total pressures for Run 40(Part 619), leeward (inboard) inlet ---------- 180


histories of engine face total pressures for Run 18(Part 624), blastward (outboard) inlet -------- 188


histories of engine face total pressures for Run 18(Part 624), leeward (inboard) inlet ---------- 196


histories of engine face distortion for Run 39(Part 544), blastward (outboard) inlet -------- 204

5.17 Comparison of theoretical and experimental timehistories of engine face distortion for Run 39(Part 544), leeward (inboard) inlet ---------- 210

5.18 Comparison of theoretical and experimental timehistories of engine face distortion for Run 40

(Part 619), blastward (outboard) inlet -------- 216

5.19 Comparison of theoretical and experimental timehistories of engine face distortion for Run 40(Part 619), leeward (inboard) inlet ---------- 222


histories of engine face distortion for Run 18(Part 624), blastward (outboard) inlet ------- 228

5.21 Comparison of theoretical and experimental timehistories of engine face distortion for Run 18(Part 624), leeward (inboard) inlet- --------- 234

7.1 Boundary layer separation and distortion from fan

reflected shock wave ----------------- 246

12

LIST OF ILLUSTRATIONS (CONCLUDED)

Figure Page

7.2 Predicted free-stream properties and form factor Hiof boundary layer on cowl of blastward inlet. Afterreflected wave enters inlet -------------- 247

7.3 Predicted free-stream properties and form factor Hiof boundary layer on ramp of blastward inlet. Afterreflected wave enters inlet -------------- 248

7.4 Predicted free-stream properties and form factor Hiof boundary layer on cowl of leeward inlet. Afterreflected wave enters inlet- ------------- 249

13

LIST OF TABLES

Table Page

1.1 16T wind tunnel test conditions ----------- 17

4.1 Post-test assessment of possible blast-inducedseparation within leeward inlet ----------- 122

5.1 Test conditions for correlation runs --------- 140

A.1 Ramp/cowl transducer evaluation ----------- 255

14

SECTION 1

INTRODUCTION

The problem of the interaction of a blast wave from a nuclear

explosion with an aircraft engine-inlet system is of importance for

military survivability/vulnerability evaluations.

An experimental study regarding this problem was recently

conducted at the Arnold Engineering Development Center (AEDC) (References 1

and 2). In this study blast waves produced by shock tubes impinged on

a 0.1-scale B-1 inlet pair mounted in the AEDC 16T transonic wind tunnel.

Tests were performed at tunnel (pre-blast) Mach numbers of 0, 0.55,

0.70, 0.85 and 0.90 for blast overpressures (scaled to 1 atm. ambient

pressure) from 2 to 6 psi for inlet flow rates representative of cruise

and maximum power conditions. These tests are described in detail in

Reference 1, which also presents a preliminary analysis of the test

data and a preliminary correlation of the test results with predictions

of the Blast Induced Distortion BID-2 computer code (Reference 3).

The present report is a continuation of the studies of

Reference 1 covering the following topics. Section 2 discusses the

blast response of the mean total pressure at the engine face. Section 3

discusses general features of blast and operational effects on engine-

face mean total pressure signature. Section 4 discusses specific

effects of blast and operational variables on engine-face mean total

pressure. Section 5 presents detailed comparisons of theoretical and

experimental inlet pressures and distortion parameters. Section 6

presents an evaluation of some large late-time inlet distortion data

discussed in Reference 1. Section 7 discusses engine reflected shock-

boundary layer interaction. Conclusions are presented in Section 8.

Since this report is closely related to Reference 1, it is

assumed herein, to limit repetition of material from that reference,

that the reader is familiar with Reference 1, particularly regarding

test geometry and test instrumentation. However, for the reader's

15

convenience, Table 1.1 and Figures 1.1 and 1.2 indicating test

conditions and pressure transducer locations in the engine inlets and at

the engine face are repeated here from Reference 1. (See Reference 1

for an explanation of the table and figures.)

16

TABLE 1. 1

16T WIND TUNNEL TEST CONDITIONS

Pirt Tube Shock Intercept YawPoInt Mach Flow Rate (lb/sec) Tube rressure Overpressure Angle Ang.

Run No.8 Inl IB Inlet No. (psia) (pSI) (deg) (deg)

1 501.01 0 0 0 2 69 2.7 79 0

2 615.03 .552 235 235 3 186 4.7 91

3 591.03 .550 351 348 1 157 3.7 764 589.03 .551 351 348 2 115 3.8 975 590.02 .549 351 349 3 124 4.0 94

6 602.02 .700 302 300 1 72 2.6 86

7 600.04 .70i 302 302 2 58 2.6 1068 573.04 .700 304 302 112 5.0 98

9 601.03 .701 302 300 3 69 3.0 10410 574.03 .701 303 300 132 4.4 97

11 621.03 .700 351 351 1 73 3.0 8412 519.02 .699 348 344 103 3.8 8813 527.02 .700 349 344 135 4.8 7814 626.02 .701 352 352 142 4.8 79

15 512.03 .700 351 344 2 59 2.8 10316 517.02 .700 349 344 | 85 3.8 10317 525.02 .701 348 344 113 5.0 10018 624.02 .700 350 350 139 '.2 99

19 513.03 .700 351 344 3 70 3.0 102

20 518.02 .700 348 343 102 4.2 9921 526.02 .700 349 344 133 4.8 9822 625.02 .701 350 350 155 5.6 92

23 570.03 .699 351 350 1 144 3.6 87 +5.024 568.04 .70(0 350 349 2 132 5.8 103 +5.025 509.03 .703 350 349 3 143 4.2 105 -5.0

26 559.02 .850 300 298 1 61 2.2 89 027 598.03 848 299 299 90 3.0 8528 584.03 .847 294 293 122 5.0 8229 608.04 .850 299 299 142 4.4 82

311 557.04 .850 300 298 2 55 - -31 596.05 .848 300 300 73 3.8 10832 582.03 .847 300 297 94 4.4 105

31 606.03 .849 300 299 120

34 558.03 .850 303 301 3 60 >2 -

35 597.03 .848 300 300 85 4.0 10436 583.03 .847 298 296 114 4.4 10237 607.03 .850 299 298 140 4.8 108

38 546.02 .847 348 347 1 121 3.6 84

39 544.04 .847 348 347 2 94 4.0 10740 619.02 .850 352 351 120 5.8 110

41 545.03 .847 348 347 3 113 4.4 10542 620.02 .850 351 351) + 139 5.6 100

43 553.03 .900 327 329 1 117 3.0 8644 550.02 .899 349 354 2 86 4.0 10745 551.01 .900 349 354 3 104 4.2 105

17

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22.5 22.5 DYNAMIC AND TOTAL PRESSURE PROBE

2 8. 2605 28o1 180

, ,..0', 28 260, _Z .- 450 ,81h, 0 /?- 89 2 80A.- 183 160 1803 65]Wb2838 M, 2603 .1o9

1 12 38 183 ,,' 1603 -,,, 65"283T 26 26 161

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283 6- 2631 18330 632 I" 1V3Laos

283. / 2)'"1

2 611 2 16.7.,1 2621 ' / 4 D 2612 16271 AU2 / 1612 /266211 -. 261 61 16284 - 1826 1 1613 q

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6 84 2*21"' 1624 82 16197 TEADY262,12257 PROBE 1625 1 25 1819k STATE

62 NUMBERS 12 TT

/LOOKING DOWNSTREAM PROBENUMBERS

INBOARD ENG~INE FACE RAKES OUTBOARDYj=0.777

* 2601 - 264.0 STEADY STATE y2z LO2S e 1601 - 164~0 STEADY STATE*@2801 - 284+0 DYNAMiIC Y3 1.393 .1801 - 184.0 DYNAMIC

.Y4: .683

Y6 : 2.149

Yy: 2.250

Figure 1.2. Engine face transducer locations. (Rockwell drawing)

19

L . ..... ..._.,.,.. . ...... ......, ... -.,,, - .... .. ..L.... ,b ;.d ,,-i

SECTION 2

RESPONSE OF MEAN TOTAL PRESSURE AT ENGINE FACE

The effect of the blast wave interaction upon the mean total

pressure at the engine face as measured in the tests is examined in

this section and the following two sections. In this section the test

records of the mean total pressures at the two engine faces are examined

qualitatively for each run on a run-by-run basis. In Section 3 the

records are grouped so as to show the effect of each test variable, one

at a time - snock overpressure, intercept angle, inlet flow rate and

free-stream Mach number - and the general effects of the variables are

examined qualitatively. In Section 4 the effect of these test variables

upon specific features of the records are examined on a quantitative

basis.

2-1 GENERAL FEATURES OF EFFECT ON MEAN TOTAL PRESSURE.

The blast wave first intercepts the cowl of the blastward

inlet, then diffracts around the cowl and into the inlet. It then

reflects from the splitter between the two inlets and diffracts around

the splitter and into the leeward inlet. At each engine face of the two

inlets a series of shock waves arrives as a result of the diffractions

and multiple reflections from the duct walls.

Typical records of the reduced mean total pressures at the

engine faces of the two inlets are shown in Figure 2.1. The quantities

R20 and R21 are the ratios of the instantaneous mean total pressures at

the engine faces of the outboard and inboard inlets, respectively, to

the preblast total pressure in the wind tunnel. The mean total pressures

for each inlet were obtained from averages of the 40 pitot probe measure-

ments at each engine face. Tube 2 was fired in this run, so R20 is the

record for the blastward inlet and R21 the record for the leeward inlet;

the same would be true for a Tube 1 firing and the opposite for a Tube 3

firing.

R20 in Figure 2.1 is constant at 0.99 until about 8.83 milli-

seconds (ms). It then jumps up to about 1.1 on shock arrival at the

engine face, followed by a steep ramp-up of about 0.33 ms duration to a

20

peak of about 1.4. The ramp is believed to consist generally of a

staircase of weak shocks produced by the multiple reflections of the

initial blast shock from the splitter and the walls of the inlet. R20

then drops off from 1.4 down to about 1.3 due to the weakening of the

upstream reflection by diffraction of the incident shock around the

splitter and other outside surfaces of the inlets.

At 9.7 ms R20 increases abruptly again due to reflection of

the shock wave from the downstream throttle formed by the control vanes.

This wave is an upstream facing shock followed by a compression wave.

The rate of rise of the engine-face total pressure varies with test

conditions and the particular characteristics of the initial shock and

compression wave within the inlet.

The strength of the upstream-facing shock/compression wave in

terms of pressure would be greater than might appear from records of

total pressure. This reflected wave slows down the flow passing through

it. The velocity reduction across the shock itself decreases the total

pressure, offsetting partially the increase in total pressure due to

pressure rise.

The reflection from a fan stage would be expected to produce

a stronger wave than occurs from the reflection from the throttle

(choked orifice). This is demonstrated by calculations for the upstream

reflection of a shock wave from a typical high-performance fan in

Reference 1, Section 8.5.

The concern regarding the upstream-facing shock/compression

wave is that it tends to distort the flow within the inlet passing

through it to an extent that stalling of the fan stages within the

engine might result. Calculations presented in Reference 1, Section 9,

demonstrated that the boundary layers on the ramp and cowl surfaces of

the blastward inlet, and possibly the cowl surface of the leeward inlet,

would separate for full-scale inlets (lOx model size) due to a repre-

sentative reflected wave (5-psi blast intercept at 90 degrees for

Mach 0.85 and 350 lb/s full-scale reduced weight flow).

21

Following the arrival at the engine face of the reflected wave

from the control-vane throttle, R20 at the engine face then decreases

slowly in this run until about 12 ms, at which time it falls off rapidly

because of the fall off of pressure within the portion of the blast wave

in the area of the inlet mouth. For firings from Tube 1, R20 often

increased at about 3 ms or so after shock arrival, due to arrival of the

cold driver gas from the shock tube.

At the engine face of the leeward inlet Figure 2.1 shows that

the inboard total pressure R21 was constant at 0.99 in the test until

9.4 ms. It then jumped to about 1.10 upon arrival of the blast shock

that had diffracted around the splitter. RTI then continued to rise

due to further blast flow around the splitter, then level off and

abruptly rose again at 10.1 ms due to the shock reflected from the

control vanes of the inboard inlet.

Beginning at about 10.4 ms R21 then fell off rapidly. This is

attributed to the large sideslip angle produced by the blast-induced

flow at the inlets, which will be discussed further. A large sideslip

angle would produce separation of the flow from the leeward surface of

the splitter or retard the flow sufficiently that it would be vulnerable

to separation by the interaction with the reflected shock from the

control vanes.

The R20 and R21 records are discussed below relative to the

effects of the tunnel Mach number, M, the inlet full-scale weight flow

at the engine face, W2R; the particular shock tube fired; and the (nominal)

overpressure of the incident blast shock, Aps.

One difference between firings from Tube 1 and those from

Tubes 2 and 3 is the blast intercept angle, *. Intercept angles from

Tube 1 ranged from 76 to 89 degrees from head-on; intercept angles from

Tubes 2 and 3 were greater, ranging from 91 to 110 degrees. The second

difference between the firings was that Tubes 1 and 2 fired from the

outboard side of the inlets and Tube 3 from the opposite side, over the

fuselage, so the blast wave from Tube 3 traveled over more of the

fuselage model.

22

2-2 MACH 0, 0 LB/S FULL-SCALE WEIGHT-FLOW.

The R20 and R21 records for Run 1 with the tunnel off and a

zero weight flow through the two inlets are shown in Figure 2.2. The

record for the blastward inlet R20 indicates that after the shock arrived

at the engine face (9.7 ms) R20 increased to about 1.26 and then fell

off with very little indication of any reflection from the control

vanes. RTO finally leveled off about 1 ms after shock arrival at about

one-half of the peak value. The abrupt rise in RTO, beginning 3.9 ms

after shock arrive (13.6 ms), is attributed to the arrival of the cold

driver gas from the shock tube, which ends the test.

The record for the leeward inlet indicates that R21 rose

stepwise at shock arrival (11.7 ms) to a level of about 1.08, continued

to climb to a maximum of about 1.14 and then decayed slowly. It is

speculated that the decay may be due to a gradual separation of the flow

from the leeward side of the splitter and ramp of the inboard inlet.

The dynamic pressure is small at Mach zero relative to the static over-

pressure so that separation would not produce as large an effect on R21

as in the example shown in Figure 2.1, where the Mach number is 0.70.

2-3 MACH 0.55, 235 LB/S FULL-SCALE WEIGHT FLOW.

Figure 2.3 shows the records for Run 2, which had the lowest

non-zero weight flow for the two inlets, 235 lb/s (full scale) each.

The tunnel Mach number was also the lowest non-zero value, 0.55, and the

incident shock overpressure was high, 4.7 psi (nominal). Tube 3 was

fired, so the inboard inlet (R21)was on the blastward side.

The record for R21 in Figure 2.3 has a high first peak, which

is attributed to the relatively high shock overpressure. The reflected

shock from the throttle control vanes, which arrived about 0.7 ms after

the first shock, appears to have been the strongest here of all of the

tests, in terms of the increment in R21. The weight flow was also low

in this test, and comparisons between tests indicated that low weight

flow increased the second rise.

23

The R20 record in Figure 2.3 shows the effect of a strong

initial shock for the leeward inlet, by the abrupt rise to a value of

1.13. It then continues to rise more slowly to 1.16, fall off, then

rise again. The latter rise is attributed to a weak reflection from the

control vanes that reaches the engine face about one millisecond after

shock arrival. The R20 record then exhibits a rapid decay which is

attributed to a large sideslip angle, produced by the blast wave, that

degrades the flow uniformity entering the leeward inlet. The combination

of a low Mach number, low mass flow and strong blast wave results in a

large sideslip angle.

This is an example where the RTO fall-off for the leeward

inlet commenced slightly after the arrival of the reflected shock at the

engine face. This would be expected to be the sequence of events if

the reflected shock caused shock-boundary layer separation within the

inlet. Further information would be needed to verify that separation

actually occurred, however it does appear clear from the record that

significant flow deterioration did take place within the leeward inlet.

This strong throttle-reflected shock in the blastward inlet

and the combination of a throttle-reflected shock in the leeward inlet

followed by a rapid decay could be detrimental to the operation of

engines attached to such an inlet. The throttle-reflected shocks,

simulating engine reflected shocks, would tend to separate the flow

within the inlets, resulting in a rapid decrease in total pressure and

possible stall of a gas turbine engine.


The results of firings at Mach 0.55 and a full-scale reduced

weight flow of 350 lb/s, (Runs 3, 4 and 5) are shown in Figure 2.4.

The results for Run 5, Figure 4c, at 350 lb/s can be compared

with those for Run 2, Figure 2.3, at 235 lb/s. Both firings were from

Tube 3 with comparable shock overpressures, 4.0 and 4.7 psi, respectively.

The first peak for the blastward inlet is similar for the two weight

24

flows, but the second peak, from the throttle reflection, is much smaller

for the higher weight flow and it also is spread out more timewise.

This characteristic of weak throttle reflections is true also for the

other two firings at Mach 0.55 and 350 lb/s, Figures 2.4a and 2.4b.

For the leeward inlet at Mach 0.55, the first peak of Run 5,

Figure 2.4c, is similar to the first peak of the 300 lb/s firing,

Run 2, Figure 2.3. The throttle reflection for the higher weight flow,

Figure 2.4c, is weaker than for the lower weight flow, Figure 2.3, as it

was for the blastward inlet. The decay rate is not as steep either, as

at the lower weight flow. It is speculated that these effects may be due

to a lower blast-induced sideslip angle for Run 5 than for Run 2 because

of the higher weight flow and 4lightly lower shock overpressure.


The records of the reduced total pressures RTO and RTI at the

engine face for tests at Mach 0.70 and a full-scale reduced weight flow

of 300 lb/s, Runs 6 to 10, are presented in Figure 2.5. For the firing

from Tube 1 (Run 6, Figure 2.5a) the shapes of the records are generally

similar to those for the example discussed in Figure 2.1. Again, the

second jump, due to the reflected shock from the throttle, which simulates

the engine, is substantial for both the blastward and leeward inlets.

The total pressures both remain high for about 3-1/2 ms after shock

arrival until cold gas appears to have arrived from the shock tube (R20)

or blast decay sets in (R21).

The weaker blast firings (2.6, 3.0 psi) from Tube 2 (Run 7)

and Tube 3 (Run 9) show a marked falloff in the blastward total pressures

after the throttle reflection arrives. This fall off is believed to be

due to decay in the blast wave arriving at the inlet. Therefore the

falloff for the leeward inlet is not attributed to sideslip angle in

this case. For the strong shocks (5.0 and 4.4 psi) from Tube 2, Run 8,

and Tube 3, Run 10, however, the blastward ratios held up well for 3 to

4 ms after shock arrival, so the falloff of the leeward ratios beginning

at about 1 ms after shock arrival is attributed to the large sideslip

angle produced by the blast wave (and not to decay in the blast properties

arriving at the inlet).

25


The results for Mach 0.70, and a full-scale weight flow of

350 ib/s, Runs 11 to 22, are shown in Figure 2.6. For Tube 1 firings,

Figure 2.6a-d, the effect of weight flow on the interaction is found by

comparison with Run 6, Figure 2.5a, for the same Mach number but only

302 lb/s. For the blastward inlet the first peak in RTO for Run 11,

Figure 2.6a, which has a nearly equal shock overpressure (3.0 psi) to

Run 6 (2.6 psi), is similar to the peak for Run 6. The second peak in

comparison to the first, is much smaller for Run 11, than at the lower

weight flow of Run 6.

The total pressures for Tube 1 at 350 lb/s hold up in both the

blastward and leeward inlets for the full 3-ms nominal test period for

the higher overpressures of 3.8 to 4.8 psi (Runs 12-14). This means

that sideslip is not causing a problem to the leeward inlet. There is a

small falloff for 3.0 psi (Run 11).

There is a feature of the blast wave that should be noted,

for later discussion, from the results shown in Figure 2.6. It will be

discussed with respect to Run 12, Figure 2.6b. The blast shock front

arrives at the blastward inlet first, of course, which is why R20 rises

about 1/2 ms before R21 for Tube 1. The interior flow of the blast wave

would also arrive first at the blastward inlet. However, the rapid

falloff in the total pressure for the blastward inlet (RTO) at about

12-1/2 ms is preceded by the falloff for the leeward inlet (RTI) by a

time of about 1/2 to 1 ms. It also decays earlier at the leeward inlet

for the firings at higher shock overpressures, as is evident in Figures 2.6c-d.

Examination of measurements of the blast waves produced by this method

has shown that this earlier decay at the leeward inlet is a result of

the three-dimensional (vs. one-dimensional) character of the blast wave.

In this case a decaying portion of the blast wave reaches the leeward

inlet before a decaying portion reaches the blastward inlet.

26

The question then arises whether the decay in total pressure

observed sometimes for the leeward inlet, which is attributed in some

cases to blast-induced sideslip, might in fact be due to this decay in

the blast pressure instead. The question cannot be resolved with

certainty at this time. But, from an examination of trends with varying

blast strengths and from measurements with external claw probes, it

appears at present that the blast decay results in a definite signature,

such as appears in Run 12, Figure 2.6b.

Another feature of the blast signature is shown in Figure 2.6c

for Run 13. At about 13 ms, R20 rises at an increasing rate. This rise

is associated with the arrival of the cold driver gas from the shock

tube. It is also observed for the leeward inlet by the rise in R21

beginning at about the same time. This rise is followed by the pressure

decay of the blast wave cited above for the leeward inlet (R21), starting

at about 14-1/2 ms, and for the blastward inlet (R20), starting somewhere

between about 15 to 17 ms. The results for Run 14, Figure 2.6d, are

similar.

These records all indicate that there is a usually good 3 ms

or more of good test flow within the blast wave before these blast

anomalies affect the results for this model. The anomalies appear to

produce characteristic signatures in the records that make them rather

well defined.

Tube 2 firings at Mach 0.70 and 350 lb/s, Figures 2.6e-h, and

Tube 3 firings, Figures 2.6i-1, produce weak throttle reflections for

the blastward inlet until the blast shock overpressure approaches about

5 psi, Figures 2.6h and 2.6k-1. For the leeward inlet there is essentially

no throttle reflection. There is however a decay for the leeward inlet,

beginning about 1-2 ms after blast shock arrival, that is attributed to

blast-induced sideslip angle. The rate of the decay increases with

blast strength which follows the trend expected for an increasing angle

of sideslip, associated with a stronger blast wave.

27

2-7 MACH 0.70, 350 LB/S, ±50 YAW.

The results with yaw angles of ±5 degrees are presented in

Figure 2.7. In each test the model was yawed with the nose away from

the particular shock tube that was fired. This means that the sideslip

angle is increased by the blast wave.

The test conditions of Run 23, Figure 2.7a, compare otherwise

most nearly with those of Run 12, Figure 2.6b. The results for the

blastward inlet are rather similar. The leeward inlet shows a notice-

ably greater falloff beginning about 1 ms after blast arrival (or at

10.5 ms). This means that the 3.6-psi blast wave with the 5-deg yaw

angle evidently produces too much sideslip for the flow within the

leeward inlet to maintain total pressure.

The test conditions for Run 24 with Tube 2, Figure 2.7b,

compare most nearly to those for Run 18, Figure 2.7h, except for the

initial yaw angle. For the blastward inlet the throttle reflection is

weaker, but otherwise the results are similar. For the leeward inlet

the falloff due to yaw is much more marked than for Tube 1, Figure 2.7a,

beginning very shortly after shock arrival.

The results for Run 25 with a firing from Tube 3, Figure 2.7c,

are similar to those for Tube 2, Run 24. The test conditions are the

same as for Run 20, Figure 2.6j, except for the -5-deg yaw. The blast-

ward inlet results are about the same as for Tubes 1 and 2, Figures 2 .7a

and b. The leeward inlet has a rapid pressure felloff, as for Tube 2.

2-8 MACH 0.85, 300 LB/S FULL-SCALE WEIGHT-FLOW.

The total pressure results at the engine face for Mach 0.85

and a weight flow of 300 lb/s are shown in Figure 2.8. Firings with

Tube I are presented in Figures 2.8a-d. The jumps due to throttle

reflections (engine simulation) for the blastward inlet are stretched

out much more in time for the weaker shocks, below about 4 psi, than

they were at the lower Mach numbers. The reflected wave steepens up

as the blast strength is increased, so that at 4.4 psi for Run 29

(Figure 2.8d) it is quite steep. The jump to the second peak is also

large.

28

The total pressures in Figures 2.8a-d for the leeward inlet

with Tube 1 firings build up well and generally hold up for the blast

period of about 3 ms. The only particular falloff occurs with the

weaker shocks of 2.2 and 3.0 psi (Figures 2.8a and b) beginning about

1.5 ms after shock arrival, which is attributed to decay in the blast

wave, since it roughly follows the fall-off of the blastward total

pressure.

For Tube 2 and 3 firings at Mach 0.85, Figures 2.8e-j, the

results are also similar to the Mach-0.7 results, Figure 2.5, for the

same tubes. The rapid falloff in the total-pressure ratios beginning

about 1-1/2 ms after shock arrival is again attributed to the decay In

the blast wave at the inlet. There is a strong effect of the throttle

reflection for the blastward inlet, also, appearing in the steepening of

the second rise and increase in level of the second peak as the blast

strength is increased. The leeward records have typical first peaks,

but again only very weak effects of throttle reflections which are

masked by a rapid decay in total pressure. There is no clear indication

of separation occurring in the leeward inlet at Mach 0.85 as there was

for the higher shock overpressures at Mach 0.70 (Figures 2.5c and e).


The results for Mach 0.85 and a full-scale reduced weight flow

of 350 lb/s are shown in Figure 2.9. For Tube 1, Figure 2.9a, the

blastward inlet for a 3.6-psi shock, Run 38, has a sharp initial peak

and a strong throttle reflection, similar to the results for Run 29,

Figure 2.8d, at 300 lb/s and a 4.4-psi shock overpressure. The results

for the leeward inlet are also similar to those with 300 lb/s.

For Tubes 2 and 3 at Mach 0.85, Figures 2.9b-e, the results

are similar to those for 300 lb/s, Figures 2.8e-1, including the rapid

decay attributed to the blast wave decay (not to sideslip).


The results for Mach 0.90, Figure 2.10, are similar to the

results for Mach 0.85, Figure 2.9, for both the blastward and leeward

inlets.

29

PRRT/PT M P1 TUBE PLOT573-04 0.70 10.21 2 19

1.60 Fis peFal Relecti oto

1.40 -l aeRoEDec a ributed to1.20 Rm O- i

CC" I.00 S

0.80

0.60

1.60 Reflection frov control1.40 va e throttle

1.20 - a_____ to1.0Shock possibl ;'parationC' 1 .00 - -w-h *A-t

0.80

0.60

0.00 4.00 8.00 12-00 16.00 20.00

TIME IN MILLISECONDS

Figure 2.1. Typical engine-f ace mean total pressures. Run 8.Mach 0.70, W211-300 lb/s (nom), Aps5 5.O psi, 0-98 dog.

30

PRRT/PT M PT TUBE PLOT

501.01 0.00 7.10 2 19

1.60 I1.40

1.200CC1 I00

0.80

0.60

1.6011.40

-1.20

0.60d

0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS

Figure 2.2. Engine-face mean total pressures. Run 1.Mach 0, W2R-0 lb/s (nom), Aps-2.7 psi (nom), f-79 deg.

31

PRRT/PT M PT TUBE PLOT615.03 0.55 12-33 3 19

1.60

1.40

1.20 -____

(C* 1.00

0.80

0.60

1.60

1.40 -_ _ ____ _

1.20

1.-00

0.80

0.60


Figure 2.3. Engine-face mean total pressures. Run 2.Mach 0.55, W2R-235 lb/s (nom) Aps 4.7 psi (nom),*-91 deg.

32

PRRT/PT M PT TUBE PLOT591.03 0.55 12.36 1 19

1 .60

1 .40

1.20

C' 1.00

0.80

0.601 .60

1 .40

1 .20

C* 1.00 -

0.80

0.60

(a) Run 3, Ap=3.7 psi (nom), 0=76 deg.


1.60

1 .40

1.20

C\J ___0_

0.80

1.60

1.40

1.20

C' 1.00

0.80-

0.600.00 4.00 8.00 12.00 16.00 20.00


(b) Run 4, Aps3.8 psi (nom), 0=97 deg.

Figure 2.4. Engine-face mean total pressures.

Mach 0.55, W2R-350 lb/s (nom).

33

PART/PT M PT TUBE PLOT590.02 0.55 12.30 3 19

1.60

1.40

1.20I00n, 1.00 -,_____0.60

0.60

1.60

1.40

1.20

r_ 1.004 008 0

0.00 4.00 .00 12.00 16.00 20.00TIME IN MILLISECONDS

(c) Run 5, Apsi4.0 psi (nom), *i94 deg.

Figure 2.4. Concluded.

34


1.60

1.40

1.20

C*4 1.00

0.80

0.601.60

1.40- _ _ _ __ _ _ _

1.20

1 .00

0.80

0.60

(a) Run 6, Aps=2.6 psi (nom), 4=86 deg.


1 .60

1.40 _ _ _ _ _ _ _ _

1 .20

1" .00

0.80

0.601.60

1 .40

1 .20

C= 1.00

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(b) Run 7, Ap 2.6 psi (nom), $'.106 deg.

Figure 2.5. Engine-face mean total pressures.Mach 0.70, W2R=300 lb/s (nom).

35

PRRT/PT M P1 TUBE PLOT573.04 0.70 10.21 2 19

1.60

1.40 - _ _

1.20

C' I .00)

0.80 _ _ _ _ _ _ _ _ _ _ _ _ _ _

O0.601.60-_ _ _ _ _ _ _ _

1.40-_ _ _ -- - - - _ _

1.20 -_ _ _

1.00 -_ _ _

0.80

0.60

(c) Run 8, Aps=5.0 psi (nom), 0=98 deg.


t .60

1.40

1.20

c1.000.80

0.601 .60

1 .40

1 .20

4.00 8.0


(d) Run 9, Ap=3.0 psi (nom), 0-104 deg.

Figure 2.5. Continued.

36


1.60-

1.40

1.20

(C~ 1.00

0.80

0.60

1.60

1.40___________I--

1.20

01.00 ~

0.80

0.60 ___ _

0.00 4.00 8.00 12-00 16.00 2600TIME IN MILLISECONDS

(e) Run 10, Aps."4.4 psi (noma), -97' deg.

Figure.2.5. concluded.

37

PRRT/PT M PT TUBE PLOT.621.03 0.70 10.23 1 19

1.60

1.40

1 .20A

C' 1.00

0.80 -

0.601.60

1.40

1 .20

0.60__ _ _ _ __ _ _ _



1 .60

1.40 A _

1.200-0

0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

0.60 -

1 .60

1.40

1.20 -_ _ _ _

CC' 1.00.-

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(b) Run 12, Apin3.8 psi (nom), *-88 deg.

Figure 2.6. Engine-face mean total pressures.Mach 0.70, W2R-350 lb/s (nom).

38

PPRT/PT M FT TUBE PLOT527.02 0.70 10.17 1 19

1.60 -__

1 .40

1.20

0 1.00

0.80

0.601.60

1.40

1.20 -_

o 1.00 -

0.60

0.60 - _ _ _ _ _ _ _ _

(c) Run 13, Aps=4.8 psi (nom), =78 deg.


1.60

1.40

1 .20 V

C 1.00

0.80

0.601.60

1.40

1.200-4

1.00

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(d) Run 14, 6psf 4 .8 psi (nom), fi78 deg.


39


1.60

1.40

1.20

C', 1.00

0.80

0.601 .60

1 .40

1.20

c' 1 .00 - __ ___

0.80

0.60

(e) Run 15, Ap,2.8 psi (nom), *=103 deg.

PPRT/PT M PT TUBE PLOT

517.02 0.70 10.19 2 19

1 .60

1 .40

1.20-

C* 1.00

0.80

0.601 .60

1.40

1.20 - - _ _ _ _

I-100 -

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(f) Run 16, Ap -3.8 psi (nom), 0-103 deg.


40


1.60

1.40

1.20

1.00

0.800.60

1.60

1.40 -_

1.20

0.80 ,,,,,.,0E . -

(g) Run 17, Aps=5.0 psi (nom), 4=100 deg.


1.60

1.40 A

1.20

~1.20 cr 1.00

0.80

0.60

Fiue .. otiud

1.40

1.20.. ,,,cl x

1.00 .. . ._ ,- _ ,

0.80

0.60


(h) Run 18, Apsw5.2 psi (nom), €=99 deg.

Figure 2.6. Continued.41

PRRT/PT M PT TUBE PLOT513.03 0.70 10-17 3 19

1.60

1 .40

1 .200-0

0.80

0.601.60

1.40 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

1 .20

S1.00 ____

0.80 -_ _ _ _ _ _ _ _ _ _ _ _ _ _ _

0.60 - ____

(i) Run 19, lAp =3.0 psi (nom), *-102 deg.


1 .60

1.40

1.20

t 1.00 n

0.60

t .60 ____________ ____________ ____________ __________ ____________

1.40

1.20

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(J) Run 20, Apo=4.2 psi (nom), 0-99 deg.


42

PART/PT M PT TUBE PLOT

526.02 0.70 10.20 3 19

1.60-

1.40 -_ _ _ _- _ _ _ _

1.20 - _ _ _ _ _ _ _ _ _ _

0:I-100

0.80

0.601.60 - _ _ _ _ _ _ _

1 .40

1.2011.-0 0

0.80 - _ _ _ _ _ _ _ _ _ _ _

0.60 _

(k) Run 21, Ap5=4.8 psi (nom), *=98 deg.


1.60

1.40-- _ _ __ _ _ _ _ _

1.200:

0.601.60

1.40

1 .20

C~ 1.00

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


(t) Run 22, Ap in5.6 psi (nom), *-92 deg.


43


1.60

1.40

1.20

0.80

0.601 .60

1 .40

1 .20

C' 1.00 -- ~-_ _ _ _

0.60 -(a) Run 23, yaw angle =+5.0 deg, Ap-=3.6 psi (nom), 0=.87 deg.


568.04 0.70 10.21 2 19

1 .60

1.40 -__ ____

1.20

S1.00 -

0.80 ____

0.601 .60

1.40

1 .20

C" 1.00

0.80 -_ _ _ _ _ _ _ _ _ _ _

0.600.00 4.00 8.00 12.00 16.00 20.00


(b) Run 24, yaw angle =+5.0 deg, Ap W5.8 psi (nom), 0-103 deg.

Figure 2.7. Engine-face mean total pressures.Mach 0. 70, W2R-350 lb/ (nom) .

44


1 .60

1.40

1.20C:)

, 1.00 -_

0.80

0.60

1.60

1 .40

1.20

S1.00 _ _ _ - -

0.80

0.60


(c) Run 25, yaw angle = -5.0 deg, Aps-4.2 psi (nom), 0-105 deg.


45

PRRT/PT. M PT TUBE PLOT559.02 0.85 11.78 1 19

1.60

1.40

1.20100

1 .60

1.40 -____

1 .20

I-100

0.80

0.60


PPRT/PT M PT TUBE PLOT598.03 0.85 11.78 1 19

1.60

1.40

CJ1.0

0.80

0.601.60r 1.40

S1.0

0.600.00 4.00 8.00 12.00 16.00 20.00


(b) Run 27, Ap5'3.0 psi (nom), 0-.85 deg.



46


1.60

1.40

0 1.200-0

0.80

0.601.60

1 .40

1.20cc~ 1.00-

0.80

(c) Run 28, Ap,=5.O psi (nom), 0=82 deg.


1 .60

1 .40

1 .20 _____0C'4 I00

0.80

0.60 -

1.60

1.40____ _

1.20

C~ .0 -o

0.80 I I0.60 I


(d) Run 29, Ap-.4.4 psi (nom), 0-82 deg.

Figure 2.8. (Continued.)

47


1.60 !

1.40

1.20

1 .410

1 .21 t.00

0.80

0.60t .60

1.40

1.20

0

n,- 1.00

0.80

0.60

(e) Run 31, Aps=3.8 psi (nom), 0=108 deg.

PR::RT/P T M PT TUBE PLOT582.03 0.85 1178 2 19

1.60

1.40

1.20

N 1.00 -

0.80

0.601.60

1.40

1.20

0.80

0 .60


(f) Run 32, APsW4.4 psi (nom), 0-105 deg.


48

.. -i i ii i i lm i i i ii i i . . . . . . . . . l4i i " . . . . . . . .] i


1 .60

1 .40

1 .20

I-100 -

0.80

0.601 .60

1.40 _ _ _ _ _ _ _ _ _ _ _

1 .20

C~ 1.00~ _0.60

(g) Run 34, Aps=2 + psi (nom).

PARI/PT M PT TUBE PLOT

597.03 0.85 11.78 3 19

1 .60

1.20------------

Of 1.00

0.601.60 -

1.40i

1.20

Of 1.00

0.80 -- -

0.600.00 4.00 8.00 12.00 16.00 20.00

TIME IN MILLISECO NOS

1h) Run, 35, Aps 4.0 psi (nom), 0-104 deg.


49


1.60- _ _ _ _ .__ _ _ _

1.40 --- 4----1.20

1.00

0.80 -- - -- t- ---- --.

0.60 -

1.60 -

.40 -- ~-- ---- I------- --- 4- 1.20 ----- I- ----- ~--

0.60

(i) Run 36, Ap =4.4 psi (nom), =1O2 deg.


1 .60

1.40

1.200:C' 1.00

0.80

0.60 __ _ _ _ _ _ __ _ _ _ _ _

1 .60

1.40 _ _ _ __ _ _ _ _ _ _ _ _ _ _I

1.20

I-100

0.80

0.60


(J) Run 37. Ap=4.8 psi (nom), =108 deg.


50

PRRT/PI M PT TUBE PLOT546.02 0.85 11.76 1 19

1.60

1.40 - ~ - ~ I-1.20

I-100 -- _-

0.60

1.60

1.40H z i -- _ - _

-1.20 tof 1.00 _ _ _ _ _ _ _ _ _

0.80- - - - - - - - - - - _ -

0.60

(a) Run 38, Ap,=3.6 psi (nom), c =84 deg.


544.04 0.85 11.76 2 19

1 .60

1.40

1.20

1. I00

0.80

1.60

1.40

1.204_

0.60


(b) Run 39, L6p =4.0 psi (nom), 0=107 deg.



51


1 .60 _ _ _ _ _

1.40

1.20

0,j I0

0' .0 ---

0.60 _____

1 .60

1.40

1.20

S1.00

0.80

0.60

(c) Run 40, Ap=5.8 psi (nom), 4=110 deg.

PRRI/PT M PT TUBE PLOT545.03 0.85 11-78 3 19

1.60

1.40

1.20 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

W" 1.00 C -___

0.601.60

1.40

- 1.20 11 .00 . ~

0.80

0.0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS

(d) Run 41, Aps4.4 psi (nom), 0-105 deg.

Figure 2.9. continued.

52


1.60

1.40

1.200-0

0.80

0.60

1.60

1 .40

1 .20

I-100 -- ze

0.80

0.60


(e) Run 42, Ap,=5.6 psi (nom), 0=100 deg.


53


1.60-

1 .40

1.201

_ 1 .000.80 ___ {j__ ___-

0.60 -_ _ _ __ _ _ _ _ _ _ _ _

1.60 - _ _

1.20 ~~1 _ I - __

0.80- -j-__0.60 h~ N

(a) Run 43, Ap,=3.0 psi (nom), 0=86 deg.4

PART/PT M P[ TUBE PLOT550.02 0.90 12.42 2 19

1.60 -_ _ _ _ _ _ _ _ _ _ _ _

1.40__ - 1CD 1.20 -- - - - -- - - - _ _ _

O 1.00 --

0.80 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _- _ _ _ _

0.60 -_ _ _ _ _____

0.60 _ _ _ _ I


(b) Run 44, Aps=4.0 psi (nom), =107 deg.

Figure 2.10. Engine-face mean total pressures.Mach 0.90, W2R=350 lb/s (nom).

54


1.60 - _ _ _ _ _ _ _ __ _ _ _

1.40 - _ _ _ _ _ _ _ _ _ _ _ _

1.20-- _ _ _ _ _ _ _ _ _ _ _ _ _

0-o

0.80 -_ _ __ _ _ _ _ '

0.60

1.60

1.40 -- - -_ _

- 1.20 ~-

0.80 - 1~0.60

0.00 4.00 -8.00 12.00 16.00 20.00TIME IN MILLISECONDS I

(c) Run 45 Aps=4.2 psi (nom), 0=105 deg.


55

SECTION 3

GENERAL FEATURES OF BLAST AND OPERATIONAL EFFECTS

ON ENGINE-FACE MEAN TOTAL-PRESSURE SIGNATURE

The effect of the blast parameters and operational parameters

on the mean total pressure at the engine face is examined in this

section. The blast parameters are the shock overpressure of the inci-

dent shock, Aps, and its intercept angle, , to the inlet. The opera-

tional parameters are the inlet full-scale reduced weight flow, W2R,

and wind tunnel Mach number, M.

The records of the reduced mean total pressure at the engine

face are grouped in Figures 3.1 to 3.56 according to test conditions

(Aps' ,, W2R, M) to show the effect of each variable separately. The

results are presented separately for the blastward and leeward inlets.

3-1 EFFECT OF BLAST SHOCK OVERPRESSURE.

The effect of the blast shock overpressure, Aps, on the RTI

and RTO records is shown in Figures 3.1 to 3.7 for the blastward inlet

and Figures 3.8 to 3.14 for the leeward inlet.

For the blastward inlet, the effect of increasing the shock

overpressure is to increase the magnitude of the first peak and to make

the shape of the peak a sharper spike as for example Figure 3.1. The

magnitude of the first peak will be examined quantitatively in Section 4.

The second peak, which is produced by the reflection from the control

vane throttle simulating the engine, steepens up in rise rate with

greater shock overpressure. The height of the second peak, referenced

to its initial point of rise, essentially does not vary with shock

overpressure.

The variation of the engine-face total pressure following the

second peak is generally independent of the shock overpressure until

the test is terminated by the arrival of the cold driver gas from the

shock tube. The cold gas appears to arrive in Figure 3.1b at about

56

11.7 ms (3.4 ms after shock arrival) and in Figure 3.1c at about 13.2 ms

(3.2 ms after shock arrival). The most obvious exception to this

independence of shock overpressure is shown in Figure 3.5 at Mach 0.70,

350 lb/s full-scale weight flow for firings from Shock Tube 3. This

decay in RTI following the second peak for the two lowest overpressuretests, Runs 19 and 20, is attributed to a decay in the blast overpressure.

The mean engine-face total pressure records for the leeward

inlet, Figures 3.8 to 3.14, have a characteristically different signature

from those for the blastward inlet. These records for the leeward inlet

show that the initial shock jump is followed generally by a continuing

rise. For example, in Figure 3.8a the total pressure following shock

arrival rises to a peak in-about 1 1/2 ms. The first part of the rise,

which reaches a plateau in about 1/4 ms, is attributed to post diffraction

effects of the blast wave around the splitter. The climb from the

plateau to the peak is attributed to reflection effects from the choked

throttle, that simulates the engine.

The magnitude of the initial (shock) jump in reduced total

pressure for the leeward inlet, Figures 3.8-3.14 increases with increasing

shock overpressure. For the seven firings from Shock Tube 1, Figures 3.8-3.9,

the shape of the pressure records for the 3-ms test period is pretty

much the same. The wave reflected from the throttle always has the

relatively iow rise characteristic for the leeward inlet.

The leeward inlet records for firings from Shock Tube 2,

Figu'es 3.10-3.11, are similar to those for Shock Tube 1 except that the

mean total pressure falls off significantly after about a millisecond,

depending upon the shock overpressure. This fall off is attributed to

the development of large flow nonuniformity and possible separation

within the inlet due to the large blast-induced sideslip angles. For

Shock Tube 3, Figures 3.12-3.14, the mean total pressure also falls off

rapidly for shock overpressures of 4 psi or more, but in these cases the

blast overpressure at the inlet exhibits a similar decay, so the

possibility of reparation being present cannot be assessed.

57

3-2 EFFECT OF BLAST INTERCEPT ANGLE.

The effect of the blast shock intercept angle, 0, on the RTI

and RTO records is presented in Figures 3.15-3.29. Firings from Shock

Tube 1 had the smallest intercept angles (most head-on). Firings from

Shock Tubes 2 and 3 produced intercepts with about the same angles

(from head-on), except that Shock Tube 3 fired from the opposite side of

the aircraft model from the other two.

For the blastward inlet, Figures 3.15-3.22, the first peak in

the reduced mean total pressure has a sharper spike for the more head-on

firings (Shock Tube 1). The second peak, from the reflection from the

throttle (simulated engine), in all cases appears to be larger for more

head-on interception. From firings from Shock Tubes 2 and 3, a signifi-

cant fall-off in mean total pressure is sometimes evident after about

one millisecond (Figures 3.16, 3.17, 3.19, 3.21, 3.22), but in all cases

the blast overpressure also decays similarly, so the decay cannot be

attributed simply to the intercept angle.

For the leeward inlet, Figures 3.23-3.29, the initial shock

jump (the first vertical rise) is generally greater for the smaller

intercept angle (Tube 1). Not much can be said about the effect of the

intercept angle for later times here because of blast decay at the inlet

for Tubes 2 and 3 in these runs, with the exception of Runs 17 and 21,

Figure 3.25. In the latter cases (Mach 0.70, 350 lb/s, 4.8-5.0 psi.)

there is a very marked effect of the intercept angle on the mean total

pressure, which is attributed to separation at the large blast-induced

angle of sideslip that results.

3-3 EFFECT OF WEIGHT FLOW.

The effect of the full-scale weight flow, W2R, on the RTI and

RTO records is shown in Figures 3.30 to 3.37 for the blastward inlet and

in Figures 3.38 to 3.45 for the leeward inlet.

For the blastward inlet the higher weight flow results some-

times in a sharper spiked first peak and nearly always in a weaker

second rise, from the throttle reflection. The magnitude of the first

58

peak does not appear to correlate well with flow rate, but it generally

decays further before the apparent arrival of the throttle-reflected

wave.

For the leeward inlet, Figures 3.38-3.45, the initial shock

jump is frequently somewhat greater for the higher weight flow. The

peak of the reflected wave, in the few cases where it can be observed,

because of the absence of the effect of rapid blast decay at the inlet,

Figures 3.38, 3.39, it is essentially independent of the flow rate. For

one case, Figure 3.41 (Mach 0.70, Shock Tube 2, Aps=5.0 psi), the flow

rather clearly appears to show the effects of separation within the

inlet. The case with the greater flow rate (Run 17) has a less evident

and weaker throttle reflection and a fall off that starts sooner.

3-4 EFFECT OF MACH NUMBER.

The effect of Mach number, M, on the RTI and RTO records is

presented in Figures 3.46 to 3.51 for the blastward inlet and Figures 3.52

to 3.56 for the leeward inlet.

For the blas'ward inlet, there is no significant effect of

Mach number on the first peak, either during the rise or decay portion.

Regarding the later period, there were only two conditions where there

were enough runs not having a large decay in the external blast wave,

Figures 3.46 and 3.47. There is no clear effect of Mach number on the

record after the first peak. In particular, there is no regular effect

on the second peak, from the throttle reflection.

For the leeward inlet, there is no apparent effect of Mach

number on the first peak either. Again, only two conditions had non-

decaying blast waves, enabling examination of the later record,

Figures 3.52 and 3.53, and neither of these show any discernable effect

of Mach number.

59

1.60

1.40

1.20

0.80

0.60

a. Run 11, Ap = 3.0 psi

1.60 -

1.40 _____1___ _

1.201

C' 1 .00 _____

0.60 -I

0.60

b. Run 12, Ap = 3.8 psi

1.60

1.40 -

1.2

1 .00 *

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


C. Run 13, xps . 4.8 psi

Figure 3.1. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 1.

60

1.60 -___ ________

1.40 -___ ___

1.20

0.80

0.60 -

a. Run 26, Ap = 2.2 psi

1.60

1.20

1~ _ _ _ _

0.80

0.60

b. Run 27, Aps = 3.0 psi

1.60

1 .20 TS1 .00

0.80

0.60

C. Run 29, Ap = 4.4 psi

1.60

1.40

1.20C0C,,jS1.00

0.80 -

0.600.00 4.00 8.00 12.00 16.00 20.00


d. Run 28, Aps 5.0 psi


61

1.60

1 .40

1 .20C

t~I .00

0.80

0.60 ___ _

a. Run 15, Ap = 2.8 psi

1 .60

1.40

1.20 - _ _

C

1 1.00 -

0.80

0.60

b. Run 16, Ap = 3.8 psi

1.60

1.40- _________ _

0.80

0.60

C. Run 17, Ap = 5.0 psi

1.60

1.40A

1.200

CC" 4.00

TIME INMLIEO S

d. Run 18, Aps . 5.2 psi

Figure 3.3. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nomn). Shock Tube 2.

62

1.60

1.40 _ _ _ _ _ __ _ _ _ _ _ _

C' 1.00 - -_ _ __ _

0.80__

0.60

a. Run 39, Aps 4.0 psi

1.60 1_ _ _ _

1.40 -I. _ _ _ _

1.20

0

1.00 . ~0.80 -

0.60

0.00 4.00 8.00 12.00 16.00 20.00TIMlE IN MILLISECONDS


Figure 3.4. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (non).. Shock Tube 2.

63

1.60

1 .40

1 .20

r_ 0 .80

0.60

a . Run 19, Aps 3.0 psi

1 .60 _____

1.40- _______ _

1.20 _______ _

I .00 -

0.60

b. Run 20, Lp = 4.2 psi

1.60 ____

1.40 j _1.201

CJ1 .00 Liz

a: 0.80 ____ _ -

0.60

C. Run 21, Ap = 4.8 psi

1 .60

1.40

1 .20

C"J 1 .000.80

0.60


d . Run 22, Aps 5.6 psi


64

1 .60

1 .40 _____

- 1.20

S1.00

0 .60 - __ __________

a. Run 34, Ap S> 2.0 psi

1 .60

1.40

C' 1.00

0.80 - - -

0.0b. Run 35, LAp 4.0 psi

1.60

1.20 -. __

- I0

0 .60I

C. Run 36, Aps 4.4 psi

1.60

1.40

1.20 ____

I .00

0.80-_____ ___

0.60

0.00 4.00 8.00 12.00 16.00 20.00


d. Run 37, Aps - 4.8 psi


65

1 .60

1 .40

1.00

0.80

0.60

a. Run 41, Aps 4.4 psi

1.60

1 .40

1.20

1 1.000.80

0.60

0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONOS

b. Run 42, Aps 5.6 psi

Figure 3.7. Effect of blast shock overpressure on engine-facemean total pressure in blastuard inlet at Mach 0.85.Weight flow 350 lb/s (em). Shock Tube 3.

66

1.60

1.40

- 1.20Cl 1.00 - __ _ _ _

0.60 I _ _ _ _ _ _ _ _ __ _ _ _

a. Run 11, Ap = 3.0 psi

1.60 1_ _ _ _ _ _ _ _ _ _ _ _

1 .40 _______ __________

1.20 ______ I_________C\j _ _ _ _ _W 1.00

0.80

0.0b. Run 12, As 38psi

1 .60

1.20

I .00

0.60 d ____ ~ _____

0.00 4.00 8.00 12.00 16.00 20.00


C. Run 13, Aps - 4.8 psi

Figure 3.8. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 1.

67

1.60

1.40

1.201.00 .. . .... - _ . . -.-

0.80

0.60

a. Run 26, APs 2.2 psi

1.601.40

1.20

o 1.00

0.800.60

b. Run 27, APs 3.0 psi

1.60

1.40

1.209-4

(%% _ _._ _ -_ _ _

S1.000.80

0.60

c. Run 29, Aps = 4.4 psi

1.60

1.40

1.00

0.80

0.6010.00 4.00 8.00 12.00 16.00 20.00


d. Run 28, Aps = 5.0 psi

Figuze 3.9. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 1.

69

1.60

1.40

1.20

I.100

0680

0 .60a . Run 15, Ap = 2. 8 psi

1 .60

1 .40

1 .20

b. Run 16, Ap, 3.8 psi

1 .60

1.40

~1.20-- _

0.80

0.60

C. Run 17, Ap 5.0 psi

1.60 j '

1.40

1 .20

1.00

0.80

0.60



Figure 3.10. Effect of blast shock overpressure on engine-facemean total pressure In leeward inlet at Mach 0.70.Weight flow 350 lb/s (nomn). Shock Tube 2.

69

1.60 __ _ _ _ __ _ _ _

1.20

0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

0 .60

a. Run 39, Apr= 4.0 psi

1.60

1 .40

1.20 _ _ _ _ _ _ _ _

c I1.00 - _ _ _ _ _

0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _

0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 40,' Aps = 5.8 psi


70

1 .60

1.4011.20 - ____ ___ ____

0.80

0.60

a. Run 19, Ap = 3.0 psi

1.60 - ____

1 .40

1.20

1 .00

0.80

0.60-___ _

b. Run 20, Ap =4.2 psi

1.60

1.40- ___ ___ _

1 .20

C' 1 .00 -___________

0.80

0.60

C. Run 21, Ap =4.8 psiS

1 .60

1.40

1.20

I 1 .00

0.80

0.60 iL0.00 4.00 8.00 12.00 16.00 20.00




71

1.601.401.20C)

, 1-00

0.80

0.60

a. Run 34, Aps > 2.0 psi

! ,601.0 -_________. _______________1.40 I1.20 --- ---.. _ -

0.80......... .......................... .......... .............

0.60

b. Run 35, Ap = 4.0 psi

1 .60

1.40 -

1.20 -

1.00

0.80

0.60

c. Run 36, Ap = 4.4 psiS

1.60

1.40

1.20C)Co: 1.00

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00




72

F ....60 t~~~~~ -' _____ ___

1.60

1.40

1.20

c 1.000.80

0.60

a. Run 41, Ap s = 4.4 psi

1.60oS

1.401.20

0.00

0.60I

0.00 4.00 8.00 1200 1600 20.0TIME IN MILLISECONDS


Figure 3.14. Effect of blast shock overpressure on engine-facemean total pressure In leeward inlet at Mach 0.85.Weight flow 350 ib/s (nom). Shock Tube 3.

73

1 .60

1.401

1.20S1.00 ____

0.80

0.60

a. Run 3, Shock Tube 1, 4)=76 deg., Ap = 3.7 psi.

1.60 -

1.40 - _ _____ ____

1.20 -_ _

0.60

b. Run 4, Shock Tube 2, 4) 97 deg., Apr 3.8 psi.

1.60

t.40

t.20

C' 1.0= -

0.80

0.600.00 4.00 8.00 12.00 16.00 20.00


c. Run 5, Shock Tube 3, 4) - -94 deg., Ap = 4.0 psi.

Figure 3.15. Effect of blast intercept angle on engine-facemean total pressure in bJlastward inlet at Mach 0.55.Weight flow 350 lb/s (nom). Blast shock over-pressure 3.8 psi (nom).

74

1.60

1.40

1.200-0cJ 1 .00

0.80 " _

0.60

a. Run 6, Shock Tube 1, 4 86 deg., Aps 2.6 psi.

1.60

1.40

1.20

S1.000.80

0.60 ,

b. Run 7, Shock Tube 2, 4 = 106 deg., APs = 2.6 psi.

1.60

1.40

1.20

:: 1.00 ,

0.60

0.60

0.00 4.00 8.00 12.00 16.00 20.00


c. Run 9, Shock Tube 3, 4 = -104 deg., Aps = 3.0 psi.

Figure 3.16. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 300 lb/s (nom). Blast shock over-pressure 2.7 psi (nom).

75

1.60

1.40

1.20

Ct 1.00 I

0.60 -

a. Run 11, Shock Tube 1, =84 deg., tAps' 3.0 psi.

1.60

1 .40

1.20

CC I .00

0.80

0.60

b. Run 15, Shock Tube 2, 4)=103 deg., Ap = 2.8 psi.

1.60

1 .40

1.20

S1.00-

0.60

0.00 4.00 8.00 12.00 18.00 20.00


C. Run 19, Shock Tube 3, 4)=-102 deg., Aps 3.0 psi.

Figure 3.17. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 2.9 psi (nain).

76

1.60

1.40-

1.201000c,,In 1- 0I _ , _ .. , _ _\

0.80

0.60

a. Run 13, Shock Tube 1, ( 78 deg., Aps = 4.8 psi.

1.60

1.40

1.20

c - 1 .00 ,_

0.80

0.60

b. Run 17, Shock Tube 2, = 100 deg., Aps = 5.0 psi.

1.60

1 .40

1.20

W~ 1.00-

0.80____

0.60


c. Run 21, Shock Tube 3, (P = -98 deg., Aps = 4.8 psi.

Figure 3.18. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 4.9 psi (nor).

77

1.60

1.40

1.20CC" 1.00

0.80

0.60

a. Run 29, Shock Tube 1, 4 = 82 deg., Aps 4.4 psi.

1.60

1.40

C3 1.20 - _ _ _

0-0

0.80

0.60

b. Run 32, Shock Tube 2, 0 = 105 deg., Ap s 4.4 psi.

1 .60

1.40ure 4.)

1.20

cc 1.00

0.80 ----

0 .60

0.00 4.00 8.00 12.00 16-00 20.00TIME IN MILLISECONDS

c. Run 36, Shock Tube 3, 0 -102 deg., Ap = 4.4 psi.

Figure 3.19. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom)j. Blast shock over-pressure 4.4 psi (nom).

78

1.60

1 .40

C3 1.20

I-100

0.80-___ _

0.60

a. Run 28, ShockTube 1, 82 deg., Ap = 5.0 psi.

1 .60

1.40

1.20

0.80-- _ __ " _- ___"

0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 37, Shock Tube 3, = 08 deg., Aps 4.8 psi.

:~;r' '.C. ff.~ct of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.85.'-"#ixht flow 300 lb/s (nom). Blast shock over-

;. ir 4.9 psi (nom).

1.60

1.40 - !1.20

CD

ctf 1.00

0.80 . .. ....

0.60

a. Run 38, Shock Tube 1, = 84 deg., Aps = 3.6 psi.

1.60

1.40 -

1.20C)

1.00 - --- _ "- -

0.80 10.60= 1

b. Run 39, Shock Tube 2, $ = 107 deg., Aps 4.0 psi.

Sf

1.60

1.40

1.20

0.80

0.60 -_"

0.00 4.00 8.00 12.00 16.00 20.00


c. Run 41, Shock Tube 3, * f -105 deg., Aps = 4.4 psi.


80

1 .60

1.40 - f1.20 .

S 1.00 - ---- -. __ _ --- _ _

0.60 V

a. Run 43, Shock Tube 1, 4=86 deg., Ap = 3.0 psi.

1 .60

t.40- _ _

CD 1.20 _ _ _

0.80 - - -- - _ _

0.60 _ _ _ _ _ _ _ _ _ _ _ _ _

b. Run 44, Shock Tube 2, 0 107 deg., Aps 4.0 psi.

1 .60 -- _ _ _ _ -1.40 - _ _ _ _ _ _ _ _

- 1.20 -_ _ _

1 .00 _ _ _ _--

0.80 K 4-

0.60

0.00 4.00 8.00 12.00 16.00 20.00TIMlE IN MILLISECONDS

c. Run 45, Shock Tube 3, =-104 deg., Aps = 4.2 psi.


81

1.60

1.40

1.20

1.00 -

0.80---

0.60

a. Run 6, Shock Tube 1, 0 = 86 deg., Aps 2.6 psi.

1.60

1.40

1.20

1.00

0.80

0.60

b. Run 7, Shock Tube 2, 4) 106 deg., APs 2.6 psi.

1.60

1.40 -

1.20

S1.00

0.80

0.60


c. Run 9, Shock Tube 3, 4 104 deg., Aps 3.0 psi.

Figure 3.23. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.70.

Weight flow 300 lb/s (nom). Blast shock over-pressure 2.7 psi (nom).

82

1.60

1 .40

1.20 -

S1.000.80- ____ ___ _

0.60

a. Run 11, Shock Tube 1, 4)=84 deg., Ap = 3.0 psi.

1.60- _______ ___ _

1.40 -____ ____ ___ _

1 .20

S1.00 -____

0.80

0.60

b. Run 15, Shock Tube 2, 4) 103 deg., Ap = 2.8 psi.

1.60

1.20 - ____ ____ ____

C)N 1 .00

0.80

0.60 - ____ ____ ____

0.00 4.00 8.00 12.00 16.00 20.00

TIMlE IN MILLISECONDS

C. Run 19, Shock Tube 3, 4)=-102 deg., Ap, 3.0 psi.

Figure 3.24. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 2.9 psi (nom).

83

1 .60

1:40 _ _ -____- _

0.60 --

0.60

a. Run 13, Shock Tube 1, =78 deg., Aps 4.8 psi.

1.60 ___ _

1.40

1 .20

0.80

0.60

b. Run 17, Shock Tube 2, 0 100 deg., *Ap 5.0 psi.

1 .60

ix00

0.04.00 8.00 12 .00 16.00 20.00


C. Run 21, Shock Tube 3, *=-98 deg., Aps 4.8 psi.


84

1.60

1.40

1.20 -_ _ _____

04 S1.00

0.80

0.60

a. Run 28, Shock Tube 1, 4-82 deg., Lp - 5.0 psi.

1 .60

1 .40

1.20 -_______ ___- ________

CD1.00

0.80 -___ _

0.600.00 4.00 8.00 12.00 16.00 20.00


b. Run 37, Shock Tube 3, *=-108 deg., Aps = 4.8 psi.


85

1.60

1.40

1.20c, 1.00 --

I=

0.80 - -i~I

0.60

a. Run 29, Shock Tube 1, € f 82 deg., Aps = 4.4 psi.

1.60

1.40

1 .20

1 .00-

0.80 "

0.60

b. Run 32, Shock Tube 2, * = 105 de., Aps = 4.4 psi.

1.60! .40 . . . .. .

1.20ot 1.00

0.80

0.60

0.00 4.00 8.00 12.00 16.00 20.00


c. Run 36, Shock Tube 3, 4 - -102 deg., Aps = 4.4 psi.


86

1.60

1.40 _ __ 1_tf 1.00 -- - .

0.800.60

a. Run 38, Shock Tube 1, ¢ = 84 deg., Aps = 3.6 psi.

1.60 -_ _ _ _

1 .40

1.20] .1 ,000.60

b. Run 39, Shock Tube 2, 0 = 107 deg., Aps = 4.0 psi.

1.60

1.40

S1.20 - *I_________ I0 .80 - ' _-_

0.80

0.60


c. Run 41, Shock Tube 3, 0 = -105 deg., Aps = 4.4 psi.


87

1.60 i_'

1 .40 i

1.20 -- -I-c1.00 -i

0:80 -____ __ ___ n_

a. Run 43, Shock Tube 1, 0 86 deg., Ap 3.0 psi.

1.60

1 .40- - - - - -_ _ _ _ _

1.20

0.60

b. Run 44, Shock Tube 2, f 107 deg., Aps = 4.0 psi.

1.60

1.40 -

1.20

0.80

0.6C

0.00 4.00 8 00 12.00 16.00 20.00


c. Run 45, Shock Tube 3, * - -105 deg., Aps - 4.2 psi.


88

1.60

1.40

1.20100

0.60

a. Run 6, W2R-FS =302 ib/s, Ap = 2.6 psi.

1.60

1.40

1.20C0(" 1 .00 .

0.80

0.60

0.00 4'.00 8.)0 12.00 16.00 20.00


b. Run 11, W2R-FS = 351 ib/s, Aps = 3.0 psi.

Figure 3.30. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 1.Shock overpressure 2.8 psi (nom).

89

1.60

1.40

1.20

t 1.00

0.80

0.60

a. Run 27, W2R-FS = 299 lb/s, Aps 3.0 psi.

1 .60

1.40

1.20 -- _°* I-0c 1.00 . . ......

0.80

0.60


b. Run 38, W2R-FS = 348 lb/s, Ap. = 3.6 psi.


90

1 .60

1.40_____

C3 1.20

" 1.00

0.80

0.60

a. Run 7, W2R-FS =302 ib/s, Aps 2.6 psi.

1 .60

1.40

1.20

S1.00 *


b. Run 15, W2R-FS = 351 ib/s, As= 2.8 psi.

Figure 3.32. Effect of weight flow on engine-face mean totalpressure in blastward inlet at M4ach 0.70. Shock Tube 2.Shock overpressure 2.7 psi (niom).

91

1 .60

1 .40

1.20C

CJ 1.00 ~

0.80

0.60


1.60 -

1.40

1 .20

0.80

0.60


b. Run 17, W2R-FS =348 ib/s, Aps = 5.0 psi.


92

7 A AB 511 KAMAN AVI TNE ULINGTON MA

F G 21/5FURTHER EVALUATI ON OF BLAST TESTS OF AN ENGINE INLET.(U)

_L MAR 79 J R RUETENIK. R F SM ILEY, M A TOMAYKO DNAOO-78-C-0239UNCLASSIFIED KA-TR-165 DNA-4994F ML

inuuuuuuuI-EEIIIIIEEEEE-EEEEEIIIEEEE-I.E.E.E..'

1.60

1 .40

1.20C0

I .00

0.80

0.60

a. Run 31, W2R-FS =300 ib/s, Ai = 3.8 psi.

1:60 Z j_1.20

1.00 =-

0.80

0.60- ____ ___ _




93

1.60

1.40

1.20

I- 100

0.8010.60


1 .60

1.40

1.20

C1. 1.00

0.80

0.60

0.00 4.00 8.00 12.00 16.00 20.00

TIME IN MI.LLISECONDS

b. Run 19, W2R-FS -344 Ib/s, Lips - 3.0 psi.


94

1.60

1.40

1.20

1.00

0.80

0.60

a. Run 10, W2R-FS = 300 ib/s, Aps = 4.4 psi.

1.60

1.40

1.20

1 1.00

0.80

0.60


b. Run 20, W2R-FS f 343 ib/s, Aps = 4.2 psi.

Figure 3.36, Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 3.Shock overpressure 4.3 psi (nom).

95

1.60

1.40

1 .00 _______

0.60

1.60

1.40 - - ___ _

1 .20 _____________I

W 1 .00 -- __ ___

0.80 -0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 41, W2R-FS - 347 ib/s, Ap8 44pi


96

1 .60

1.40

1.20

I-.00 - _ _

0.80

0.60

a. Run 6, W2R-FS =300 lb/sb A1)P 2.6 psi.

t.60r 1.401.20 ____

c~1.00 - ~ ____

0.60

0.60


b . Run 11, W2R-FS =351 ib/s, Aps = 3.0 psi.

Figure 3.38. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 1.Shock overpressure 2.8 psi (nom).

97

1.601.40

1.20

I-100

0.80

0.60

a. Run 27, W2R-FS = 299 ib/s, Aps 3.0 psi.

1.40 _ ___

1.60

1 .20 --- 1 -- -

S0.80 ' ... " ::.... '

0.60

0.00 4.00 8.00 12.00 16.00 20.00TIMF IN MILLISECONDS



98

1 .60

1.40

1.20-'-4

S1.00

0.80

0.60 ____ ____ ____ ___ _


1 .20

Cr_ 1.00

0.80

0.60 -_____

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 15, W2R-FS =344 ib/s, Al,5 = 2.8 psi..


99

1 .60

1.40 -___

1.20 __ _

S1 .00 - _ _ _

0.80

0.60


1.60

1.40 -___ ____

N 1 00 _j

0.80

0.60

I0.00 4.00 8.a0 12-00 16.00 20.00TIME IN MILLISECONDS

b. Run 17, W2R-FS - 344 lb/s, Aps = 5.0 psi.


100

1.60

1.40____ _

1.20 _ _ _ __ _

I-100-

0.80

0.60

a. Run 31, W2R-FS 300 ib/s, Ap = 3.8 psi.

1.60

1 .40 Z_

16.0 20.00

TIME Z*N MILISECNDb. Run 39, W2R-FS 347 ib/s, Aps = 4.0 psi.

Figure 3.42. Effect of weight flow on engine-face mean totalpresur inleeward inlet at Mach 0.85. Shock Tube 2.

Sokoverpressure 3.9 psi (nom).

101

1 .60

1.40

1.200-0

0.80

0.60


1.60 -_ _ ____ _

t.40

1.20

1 .00 ~

0.80

0.60

0.00 4'.00 8*.00 12.00 16.00 20.00TIME IN MILLISECONDS

b. Run 19, W2R-FS =351 ib/s, Aps 3.0 psi.


102

1.60

1.40

1.2

I-100

0.80 'I11 ____

0.60


1.60

1.40 _______ _

1.20

0.60 - ____________________




103

1.60

1.40 ___ ___ ___-

S1.0 --

0.60


1.60

1.40

1.20- -_ _- _ _ _ _

100

0.80 _______________________-

0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 41, W2R-FS = 348 ib/s, Ap = 4.4 psi.


104

1.60

1.40

1.20

I-1 00 ~

0.80

0.60

a. Run 6, Mach 0.70, Ap = 2.6 psi.

1.60

1.40 " '

1.20 A

CI 1.00

0.80

0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 27, Mach 0.85, Aps 3.0 psi.

Figure 3.46. Effect of Mach number on engine-face mean totalpr ssure in blastward inlet Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure2.8 psi (nom).

105

ism

1.60

1.40

C3 1.20Vck 1.00 -_ _ _ _ _ _ _ _

0.60

a. Run 3, Mach 0.515, Ap, 3.7 psi.

1 .60

1.40- _____ _

1.20

1 1.00

0.80 -___

0.60


1 .60

1.40

1.20

0.60

0.00 4.00 8. *00 12.00 16.00 20.00TIME IN MILLISECONDS

C. Run 38, Mach 0.85, Ap = 3.6 psi.

Figure 3.47. Effect of Mach number on engine-face mean totalpressure in blastward Inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom).

106

1.60

1.40

1.20

1 1.00

0.80

0.60

a. Run 4, Mach 0.55, Aps = 3.8 psi.

1.60

1.40

1 .0

0.80

0.60

b. Run 16, Mach 0.70, AP = 3.8 psi.

1.60

1.401-I ___

1.20 r

.00-

0.80 _

0.60

c. Run 39, Mach 0.85, LPs = 4.0 psi.

1.60

1.40 - - -- __ _

1.20

Ck 1 00 .. . . . . .. -

0.60 -___'__

0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDSd. Run 44, Ilach 0.90, Ap 4.0 psi.

Figure 3.48. Effect of Mach number on engfne-face mean totalpressure in blastward inlet. Weight flow 350 lb/s

(nom). Shock Tube 2. Blast shock overpressure

3.9 psi (nom).

107

1.60

1.40

1.20 A(D

o 1.00

0.80

0.60

a. Run 16, Mach 0.70, Aps = 3.8 psi.

1.60 _

C)1.20 /%

S1.00 -

0.80 __ _ __ _ __ _ __ _ __ _

0.600.00 4.00 8.00 12.00 16.00 20.00

TIME IN MILLISECON5

b. Run 39, Mach 0.85, Aps f 4.0 psi.

Figure 3.49. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom).

108

1.60

1.00

0.80

0.60

a . Run 10, Mach 0.70, Ap = 4.4 psi.

1.60

1.40 -

1.20

CC" 1.00 I

0.80 j- -

0.60

0.00 4.00 8.010 1.2.00 16.00 20.00



Figure 3.50. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom).

109

1.60

t1.40

1 .20 A-~ 1.00100_1_1 _1

0.'80

0.60

a. Run 5, Mach 0.55, Ap - 4.0 psi.

1.60

1.40

1.20

0.80- ____ ___ ___ _

0.60

b. Run 20, Mach 0.70, Ap = 4.2 psi.

1 .60

1 .40

1.20

0.00 4.00 6.00 12-00 16.00 20.00TIME IN MILLISECONDS

C. Run 41, Mach 0.85, Aps 4.4 psi.

Figure 3.51.. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shack overpressure4.2 psi (nom).

110

1.60

1.40

1.20

S1.00

0.80

0.60

a. Run 6, Mach 0.70, Aps = 2.6 psi.

1.60

1.40

-. 1.20"

-1.00

0.80

0.60

0.00 4.00 8.00 12.00 16.00 20.00



Figure 3.52. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure

2.8 psi (nom).

- 111

1.60

1.40

1.20

0.80

0.60

a. Run 3, Mach 0.55, Aps 3.7 psi.

1.60

1.40

1.20

(C" 1.00 -_ _ _ _ _-

0.80 I

0.60 - ____ 1

b. Run 12, Mach 0.70, Ps = 3.8 psi.

1.60

1.40 I1.20

0.60


c. Run 38, Mach 0.85, Aps - 3.6 psi.

Figure 3.53. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom).

112

1 .60

1.40

1.20-40cu 1.00 ..... . .. .. ..

0.80

0.60

a. Run 16, Mach 0.70, Aps = 3.8 psi.

1.60

1.40

1.20t"'l 1.00 - ]

0.80

0.60 _.


b. Run 39, Mach 0.85, Aps = 4.0 psi.

Figure 3.54. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure

3.9 psi (nom).

113

1.60-

1.401

1.20 IC',1.0o

0.80-- 1iij ____ _

0.60

a. Run 10, Mach 0.70, LAp 4.4 psi.

1.60 __

S1.20 ___

0.80 - . I-

0.60

0.00 4.00 8.00 12.00 16.00 20.00


b. Run 36, Mach 0.85, Apr= 4.4 psi.


114

1.60

1.40 I____ _

1 .20

0.8

0.60

a. Run 5, Mach 0.55, bps= 4.0 psi.

1 .50

1.40 - _ _ ___ ____ __

1 .20

W 1.00

0.60 ___ _

0.60

b. Run 20, Mach 0.70, tAp 4.2 psi.

1.60 -____ .____

0.801 ____I

0.60 - ____ ______________


C. Run 41, Mach 0.85, bps= 4.4 psi.


115

SECTION 4

SPECIFIC EFFECTS OF BLAST AND OPERATIONAL

VARIABLES ON ENGINE-FACE MEAN TOTAL PRESSURE

In Section 3, the general features of the reduced mean total

pressure at the engine face, RTI and RTO, are examined as a function of

the test variables. In this section factors affecting the first and

second peaks for the blastward inlet, for which the first peak is the

larger, and separation within the leeward inlet are examined.

4-1 VARIATION IN FIRST PEAK OF MEAN TOTAL PRESSURE FOR BLASTWARD

INLET.

The first peak in the mean total pressure increment at the

engine face is generally during the blast period the highest. Therefore

its magnitude is an evident measure of the effect of the wave in the

inlet. The factors affecting this magnitude will be analyzed in this

section from the test results and BID code calculations.

4-1.1 Effect of Incident Total-Pressure Increment.

In Reference 2.1, Section VII, it is shown that the change in

the mean total pressure at the engine face for the blastward inlet has

about the same ratio to the incident total-pressure change (i.e.,

across the blast shock) for intercept angles of 90, 105-and 135 degrees.

This comparison was based on results of BID code calculations. In this

section the relationship will be examined for the test data.

The jump in mean total pressure at the engine face, from the

preintercept value to the first peak after shock arrival, Ap' int

2Figure 4.1, was measured for each run. It is plotted in Figure 4.2

against the jump in total pressure across the blast shock Ptos , in the

wind tunnel. The straight line through the median of these data

(50 percent on each side of the line) is given by

P2= 1.41 Apt (4.1)t2

t0

Ninety percent of the data points fall within the +11 percent of this

straight line.

116

This result means that the jump in the mean total pressure at

the engine face to the first peak was roughly 1.41 times ar large as the

jump in the total pressure across the blast shock in the free stream.

Because the spread was essentially only ±11 percent, it is concluded

that the jump at the engine face was primarily a function of the jump at

the shock.

There are two conclusions from this result. First, the blast-

ward inlet acted somewhat as an amplifier of the total pressure increase

produced by the blast shock, by a factor of approximately 1.41. Second,

the jump in total pressure across the blast shock is the principal

factor causing the jump in mean total pressure at the engine face.

Perhaps the 11-percent accuracy of prediction (90 percentile)

may be sufficient for most purposes. But an examination will be made in

the remainder of this section to identify the factors involved in this

remaining spread.

4-1.2 Effect of Test Variables on Engine-Face First-Peak Ratio.

In this subsection, the data will be examined to determine the

Ap I

dependence of __2 , called hereafter "the engine-face peak ratio,"AP

t

0

upon the test variables.

4-1.2.1 Shock Overpressure Effect.

The engine-face peak ratio is plotted in Figure 4.3 as a

function of shock overpressure for Mach 0.70 and 0.85 and full-scale

reduced weight flows of 300 and 350 lb/s. There is no apparent con-

sistent trend with shock overpressure.

4-1.2.2 Shock Intercept Angle Effect.

The variation of the engine-face peak ratio, with intercept

angle is presented in Figure 4.4 for full-scale reduced weight flows of

300 and 330-350 lb/s. Over the limited range of intercepts tested,

there is no clear variation with intercept angle at either mass flow rate.

117

Resulcs of calculations made with the BID code are plotted for

comparison in Figure 4.4b. Results for angles of 135 degrees and

greater are not plotted because the reflected wave from the throttle

returned to the engine face before the first peak was attained. The BID

results indicate the peak ratio is essentially constant over the region

of the test data. The BID value is at the low end of the test data,

which is attributed to some dissipative effects in the calculations.

The BID code results indicate that the peak ratio would be less than

the approximate value of 1.41 from these tests.

4-1.2.3 Weight Flow Effect.

The engine-face peak ratio is plotted in Figure 4.5 as a

function of the full-scale reduced weight flow, W2R.

The effect of the weight flow on the peak ratio appears to

depend upon the strength of the incident shock, based on the data for

Mach 0.70, Figure 4.5a. For shocks of 2.2 to 3.0 psi, the peak ratio

decreases with weight flow (comparative data are available only for

Mach 0.70). For stronger shocks, of 4.0 to 5.0 psi, the ratio at

Mach 0.55 decreases with weight flow, at Mach 0.70 it is constant or

increases and at Mach 0.85 it increases markedly with weight flow.

4-1.2.4 Mach Number Effect.

The engine-face peak ratio is presented in Figure 4.6 as a

fun, ion of Mach number.

There appears to be some effect of Mach number on the peak

ratio, but it is a weak effect, and the dependence appears to be a

function of the weight flow, W2R. At 300 lb/s, the trend is weak, but

it slightly decreases with Mach number. At 350 lb/s, the peak ratio

increases with Mach number.

4-2 VARIATION IN SECOND PEAK OF MEAN TOTAL PRESSURE AT ENGINE

. FACE FOR BLASTWARD INLET.

The wave entering the inlet, that produces the first pressure

peak In the reduced mean total pressure at the engine face, RTI and

RTO , Figure 4.., partially reflects from the engine (and also passes

118

through). A principal concern about the reflected wave is the possibility

of its producing separation of the flow within the inlet, by means of

the adverse pressure gradient it presents. The separation in turn could

produce unacceptable distortion at the engine face, resulting in stall

or surge and engine flameout.

The characteristics of the second peak at the engine face are

a function of the blast and inlet variables, as shown in Section 3.

These characteristics include the steepness of the rise to the second

peak, the magnitude of the second peak and the time separation between

the first and second peaks. The factors causing separation of the flow

under these transient conditions are not well understood. Therefore no

attempt will be made here to relate quantitatively the characteristics

of the second peak with the blast and inlet variables, as was done for

the first peak in Section 4.1. Instead only the qualitative characteristics

of the second peak will be reviewed here from the data presented in

Section 3.

4-2.1 Effect of Shock Overpressure.

The shape of the rise in RTI and RTO to the second peak is

found to be a function of the shock overpressure. As the overpressure

is increased, the rise steepens from a compression wave to a quite

distinct shock wave (instantaneous pressure rise). This steepening is

expected to have a significant effect on the boundary-layer separation.

This steepening of the rise to the second peak is in contrast

to the observed rise time to the first peak. The latter was essentially

independent of the shock overpressure.

The magnitude of the rise, from the initial point of rise to

the second peak, is found to be essentially independent of the shock

overpressure.

4-2.2 Effect of Intercept Angle.

The rise to the second peak is steeper to some degree for

lower intercept angles (more head-on). In general the magnitude of the

jump to the second peak also increases with decreasing intercept angle.

This is to be expected from the dependence of the first peak on the

119

intercept angle: as the angle is decreased, the jump in incident total

pressure increment increases, so the magnitude of the first peak is

greater. This stronger wave then results in a stronger reflection from

the throttle (simulated engine effect).

4-2.3 Effect of Weight Flow.

The magnitude of the rise in RTI and RTO to the second peak is

essentially always greater for the lower weight flows. This resulted

for most of the Mach numbers tested and for all three shock tubes.

This result may have importance for aircraft in flight under

high thrust conditions, such as during base escape, where the weight

flow would be high. The weaker reflected wave might result in reducing

the possibility for separating the flow within the inlet and distorting

the flow at the engine face.

4-2.4 Effect of Mach Number.

There was no observable effect of the Mach number on the

strength of the reflected shock wave, as measured by the rise to the

second peak of RTI and RTO.

4-3 SUMMARY FOR BLASTWARD INLET.

The magnitude of the jump in mean total pressure at the engine

face,AP' in these tests was primarily proportional to the jump in

total pressure across the incident shock, Aptos, within +11 percent

(90 percentile). (The remaining effect was a function principally of

the engine-face corrected weight flow and slightly a function of Mach

number.) Therefore the jump would increase with increasing shock over-

pressure and decreasing intercept angle.

The first peak of the mean total pressure at the engine face

was more peaked (sharper, spike) for intercepts at smaller angles,

more head-on.

The reflected shock from the engine-simulation throttle pro-

duced a second peak in the mean total pressure at the engine face. This

reflected wave is of concern because of possible separation of the flow

within the inlet that might result from its presence.

120

The pressure rise to the second peak steepened with increasing

shock overpressures and, to some degree, with lower intercept angles.

The magnitude of the jump at the engine face definitely increased with

lower intercept angles. Lower weight flows also produced greater rises

to the second peak, but Mach number had no apparent effect.

The BID code correlates with the trend of the variation of the

jump to the initial peak in total pressure as a function of shock

intercept angle.

4-4 FEATURES OF MEAN TOTAL PRESSURE FOR LEEWARD INLET.

The profile of the engine-face mean total pressure for the

leeward inlet varied considerably between runs. However there were

definite characteristic features of most of the profiles which are

illustrated in Figure 4.7.

The initial shock jump for the leeward inlet, as for the

blastward inlet, is one-half or less, generally, of the jump to the

peak. This initial jump is produced by the first shock that comes down

the inlet. That shock is followed by other shocks, produced by shock

reflections and diffractions, which produce the further pressure rise.

The total pressure in the records then generally levels off

until reflections of these waves return from the throttle (simulated

engine). For the remainder of the test period after the second peak

the total piessure either remains essentially steady or it falls off.

The fall-off is attributed in some cases to decay of the blast wave at

the mouth of the inlet; in other cases it is attributed to possible

separation of the flow from the inlet walls. Of all these effects,

separation is believed to be the most detrimental to engine operation,

so attention is focused here on identifying the test conditions leading

to this possible separation.

4-5 FLOW SEPARATION WITHIN LEEWARD INLET.

The assessment of blast-induced separation is tabulated for

each run in Table 4.1. The test conditions are tabulated in columns

two through five; the assessment of separation is listed in the last

three columns.

121 -; -"

TABLE 4.1

POST-TEST ASSESSMENT OF POSSIBLE BLAST-INDUCED SEPARATIONWITHIN LEEWARD INLET

Assessment ofNominal Blast-InducedFull-Scale Nominal Separation in InletReduced Shock

Nominal Weight Flow Tube Overpressure Masked ByRun Mach No. (lb/s) No (psi) None Likely Blast Decay

1 0 0 2 2.7 X

2 0.55 235 3 4.7 x

3 0.55 350 1 3.7 x4 0.55 350 1 3.8 x5 0.55 350 1 4.0 x

6 0.70 300 1 2.6 x7 0.70 300 2 2.6 x8 0.70 300 2 5.0 ?9 0.70 300 3 3.0 x

10 0.70 300 3 4.4 x

11 0.70 350 1 3.0 x12 0.70 350 1 3.8 X13 0.70 350 1 4.8 x14 0.70 350 1 4.8 X

15 0.70 350 2 2.8 x16 0.70 350 2 3.8 x17 0.70 350 2 5.0 x18 0.70 350 2 5.2 x

19 0.70 350 3 3.0 x20 0.70 350 3 4.2 x21 0.70 350 3 4.8 x22 0.70 350 3 5.6 X

23 0.70 350 1 3.6 ?24 0.70 350 2 5.8 x25 0.70 350 3 4.2 x

26 0.85 300 1 2.2 x27 0.85 300 1 3.0 X28 0.85 300 1 5.0 x29 0.85 300 1 4.4 X

122

TABLE 4.1 (Continued)

Assessment ofNominal Blast-InducedFull-Scale Nominal Separation in InletReduced Shock

Nominal Weight Flow Tube Overpressure Masked ByRun Mach No. (lb/s) No (psi) None Likely Blast Decay

30 0.85 300 2 -

31 0.85 300 2 3.8 X32 0.85 300 2 4.8 x33 0.85 300 2 -

34 0.85 300 3 >2 x35 0.85 300 3 4.0 X

36 0.85 300 3 4.4 x37 0.85 300 3 4.8 X

38 0.85 350 1. 3.6 X39 0.85 350 2 4.0 x40 0.85 350 2 5.8 x41 0.85 350 3 4.4 X42 0.85 350 3 5.6 x

43 0.90 350 1 3.0 ?

44 0.90 350 2 4.0 X45 0.90 350 -3 4.2 x

123

Separation is assessed in Table 4.1 as either: (a) none,

(b) likely or (c) masked by blast decay. "None" means that there is no

apparent sign of separation. "Likely" means that the mean total pressure

drops as qualitatively expected if separation were present within the

inlet, so it is deduced that separation was a likely cause. (This

possibility has been examined in Section 7 for one run by boundary layer

calculations.) "Masked by blast decay" indicates that separation is not

excluded, but that it cannot be verified by visual examination of the

records because of decay that was present in the blast at the inlet

mouth.

At Mach 0 and 0.55 separation appears to have occurred for all

but one case. That case was for the weakest shock of the tests at 350 lb/s

full-scale reduced weight flow.

At Mach 0.70, for the inlet unyawed, separation evidently dia

not occur for Tube 1 in any test. Separation is believed to have

occurred for Tubes 2 and 3. This would indicate that the intercept

angle is an important factor. Whether shock overpressure is also an

important factor cannot be assessed satisfactorily from these runs

because the low overpressure firings from Tubes 2 and 3 also had blast

decay present at the inlet. For the remaining tuns, Tubes 2 and I had

higher overpressures, except for one firing which was equal (4.8 psi).

For all firings with the inlet yawed, separation was either

likely or, in the case of Run 23, possible.

At Mach 0.85 there was no separation evident for Tube 1,

and blast decay masked the results for all firings from Tubes 2 and 3.

At Mach 0.90 it is possible that separation occurred for Tube 1, but

separation was masked by the present of blast decay for Tubes 2 and 3.

Blast decay at the leeward inlet was a result of the limita-

tions in the size of the blast wave that was produced by the shock

tubes. in spite of the large size of the tubes (22.6-in. ID), the

fraction of the volume of the blast wave that was satisfactory for blast

simulation was relatively small by the time the blast wave reached the

inlets. At the higher Mach numbers the wave was masked downstream

enough that the wave was unsatisfactory at the leeward inlet after a

millisecond or so. Great care was taken to locate the shock tubes to

124

meet the greatest range of test conditions within the restrictions of

the wind tunnel structure, but the blast wave at the leeward inlet could

not be maintained without sacrificing blast simulation at the blastward

inlet. The limitation in the size of blast waves that can be produced

in a wind tunnel does present a restriction to wind tunnel testing of

blast effects.

4-6 CONCLUSIONS.

The jump in mean total pressure at the engine face of the

blastward inlet following blast intercept is found to be nearly pro-

portional (+1 percent) to the jump in total pressure across the blast

shock for the range of blast conditions tested (Aps=2.2-5.8 psi,

9=76-110 deg, W2RFS=0-354 lb/s, M=0-0.90). Therefore it primarily

increases with increasing shock overpressure and decreasing intercept

angle.

The reflection of the blast wave upstream from the engine

(simulated by a choked throttle) is expected to be a principal factor to

distortion, particularly for the blastward inlet, because of shock-

boundary layer interaction, possibly resulting in separation of the flow

within the inlet from the walls. The reflection produces a second peak

in total pressure at the engine face for the blastward inlet. The rise

in total pressure approaching the second peak steepens with increasing

shock overpressure, tending to result in formation of an upstream-facing

shock within the inlet at a point nearer to the engine end of the inlet,

where the reflection enters. The magnitude of the rise to the second

peak increases for lower intercept angles (more head-on) and lower

weight flows (low engine thrust conditions). It is essentially unaffected

by Mach number.

The mean total pressure at the engine face increased more

slowly with time for the leeward inlet, following blast intercept, and

by a smaller amount.

125

A rapid fall-off in total pressure from one millisecond or

more after shock arrival that was observed in many cases for the leeward

inlet is attributed to possible separation within the inlet. It occurred

for most of the tests at Mach 0 and 0.55. At Mach 0.70 it occurred

essentially only for firings from Shock Tubes 2 and 3, which produced

higher angle intercepts. At Mach 0.85 and 0.9 it could not be assessed

because of the limited size of the blast wave.

126

First peak

Second peak

AppPt 2 it 2

Time

Pt2

RT = RT = RTI for inboard inletPt0 = RTO for outboard inlet

Apt 2

ART =

Pt 0

Figure 4,1. Sketch illustrating first and second peaks of

engine-face mean total pressure for blastward inlet.

127

0.6

Ap' AP' 1. 41 Ap 0t2 t t -

2_ 2 s

Pt0 05 -0

0 000

0

0.4 0

CD0

0

0.2

0.1

0 _0 0.1 0.2 0.3 0.4 0.5

Pt 0

Figure 4.2. Inczement in mean total pressure at blastvard engine-faceversus incident total pressure increment.

128

AP'

Ap~

s W2R-FS

2.0Tube 300 350

1 A2 0 U3 0 *

1.8

1.6

1.4 00

1.2

1.0 I0 1 2 3 4 5 6

Ap 0 (psi)S

a. Mach No. 0.70

Figure 4.3. Engine-face peak ratio versus incident overpressure.

129

2.2r

Ap't 2 I

Ip Tube W2R-FSs 2.oQ 300 r350

2 0 U

3 0 01.8 L....... _ _

1.6

@00

1.4

1.2

0 12 3 456

Ap 0 (psi)a

b. Mach No. 0.85


130

4)4

'a 00

0-C3

00

.14

0w 00 Q

0

9: 41-4

0 C

4 44

1313

.0 C-4 crr -jc0 -0

1N 00-:0 ~ ~ 0 nn 000 % 41

0 0 0; 0

C14

I C,

+0

0

-01-4

c0 c00

<0 0

Y) 0

0 0

.

414

0.3

2.2-

2.0-

t 2

6p t

S 1.8-

1.6-

Tube_2

1.4-

1.2

1.0 I

200 250 300 350 400 450W2R-FS (ib/'s)

a. M = 0.70, Ap = 2.2-3.0 psi.

Figure 4.5. Engine-face peak ratio versus weight flow.

133

2.2

2.0

t2

S

1.8

1.6

Tube 2 M 0.7

1.4 - Tube 3

1.2 Tube 3 M =0.55

1.c III200 250 300 350 400 450

W2R-FS (ib/s)

b. M - 0.55-0.70, Aps . 0-5.0 psi.


134

2.2

Tube

2.0

t 2 2Apt

0s 1.8

1.6

Tube1/2

1.4 - 3~-- y

0

1.2

1.0 1

200 250 300 350 400 450

W2R-FS (ib/s)

C. m= 0.85, ts = 3.8-4.4 psi.


135

2.2

Tube Ap

1 2.2-3.0

2.0 0 2 3.8-5.0

0 3 4.0-4.8

1.82

Ap tP0

S

1.6

Tube 1

1.4

2 -

1.2

1.0 i I I I

0.5 0.6 0.7 0.8 0.9 1.0Mo0

a. W2R-FS = 300 lb/s.

Figure 4.6. Engine-face peak ratio versus Mach number.

136

Tube 6PS0

A 1 3.7-4.2

2.0 -. 2 3.8-4.0

S 3 4.0-4.4

t2

Ap t

1.6-Tue1

1.4

1.2

1.0vII0 .5 .6 .7 .8 .9 1.0

1110

b. W2R-FS =350 lb/s.


137

Throttlereflection

Typicalwithout

Further Peak decay

Shock rise

jump

• L

4 Plateau

pN

Typical decayattributed to either

Pre-intercept separation withinlevel inlet or decay of

blast at inlet

I

Figure 4.7. Sketch illustrating features of engine-face meantotal pressure for leeward inlet.

138

SECTION 5

COMPARISON OF THEORY AND EXPERIMENT

Concurrent with the performance of the inlet blast tests of

Reference i, Kaman AviDyne developed a theoretical two-dimensional

computer code, designated BID (blast-induced-distortion), for predicting

the transient flow field produced in a inlet by a blast wave striking

the inlet at an arbitrary angle of incidence (Reference 3). Preliminary

calculations of inlet pressure time histories were made with this code

for four conditions similar to those of the experimental tests and the

results of these calculations were compared with the test results in

Reference 1. These preliminary calculations were made under the assumption

that the blast wave could be represented as having a constant orientation

and strength throughout the blast event. The code results based on

these simplifications indicated that the major features of the transient

pressures observed in the inlet tests are well represented by the BID

code results. However. a more extensive comparative study of the test

results appeared to be desirable to more definitively establish the

limitations of the code predictions. To provide such a correlation, a

detailed analysis was made of the time histories of the blast wave

strength and orientation after striking the inlet for three blast test

runs (see Table 5.1) and BID calculations of inlet response were made

using these quantities as inputs. Results of the code runs and data

correlations are presented below.

5-1 BLAST INPUT CONDITIONS.

In order to perform calculations of blast response charact-

eristics by the BID code it is necessary to represent the blast pressure,

density and velocities incident on the inlets as a time dependent plane

wave. These characteristics were determined for the three runs considered

here by a detailed synthesis of the data obtained from three claw-static

probes located around the inlet (see Reference 1), with first order

corrections being made for probe-model interference. The resulting

blast input variations for the three runs are presented in Figures 5.1

139

TABLE 5.1

TEST CONDITIONS FOR CORRELATION RUNS

Run Number 39 40 18

Part Number 544 619 624

Pre-Blast Conditions:

Mach Number 0.85 0.85 0.70

Mass Flow (ib/s) 348 352 350

Blast Conditions:

Nom. Shock Overpressure (psi) 4.0 5.8 5.2

Intercept Angle (Deg) 107 110 99

140

- -- . C'°-

through 5.3. It should be noted that the blast characteristics are

defined for a long duration of over 12 ms for the first two cases

(Run 39 (Part 544) and Run 40 (Part 619)) but only for less than 4 ms

for the third case (Run 18 (Part 624)). In the last case the test

geometry resulted in large not-readily-analyzed non-linear flow dis-

turbances at later times, associated with arrival at the inlets of the

cold air jet from the shock tube which produced the blast wave.

5-2 INLET PRESSURES.

Theoretical calculations of inlet ramp and cowl pressure time

histories from the BID code (Reference 3) are compared with the corres-

ponding AEDC test data (Reference 2) in Figures 5.4 to 5.9 for Runs 39,

40 and 18. Ordinates in these figures are the ratio of inlet static

wall pressure tc pre-blast wind tunnel total pressure and the numbers to

the left of the ordinate scales represent the transducer identification

number, transducer locations being indicated in Figure 1.1.

It is seen that the BID code results are generally in good

agreement with all major features of the test data, particularly for

Runs 39 and 40. The only apparent conspicuous differences are questions

of the relative amplitudes of the calculated and experimental pressures

for a few transducers (e.g., transducer 2902 for Run 39 in Figure 5.5)

and there is strong evidence that most of these differences can be

attributed to errors in calibration factors used to reduce the test data,

as is discussed in the Appendix.

5-3 ENGINE FACE PRESSURES.

Considering next total pressures at the engine face location,

Figures 5.10 to 5.15 present comparisons of BID code predictions with

time histories of experimental total pressures for a variety of positions

at the engine face for both the blastward (outboard) and leeward (inboard)

,

In order to permit a reasonable quantitative comparison of theory andexperiment in this report, it was necessary to take into account the factthat the pre-blast (steady-state) pressures were slightly different forthe two cases. This difference was taken into account in pressure-timeplots by vertically shifting the theoretical curves so that the pre-blaststeady-state values were the same for theory and experiment.

141

inlets. Ordinate scales are the ratio of total pressure to pre-blast*

wind tunnel total pressure and the ordinate label designates the trars-

ducer identification number, transducer locations being indicated in

Figure 1.2.

5-3.1 Blastward Pressures.

BID predictions of engine face total pressures in the blastward

inlet are compared with the corresponding AEDC test data for Runs 39,

40 and 18 in Figures 5.10, 5.12 and 5.14, respectively.

Considering first Run 39, the BID pressure variations are seen

to follow well both qualitatively and quantitatively the major features

of the test data for all of the transducer locations. To be sure, the

code results do somewhat tinderestimate the initial rate of pressure

rise, the first pressure maximum, and some very rapid shock-like changes.

Considerably smaller cell sizes would have had to be used in the BID

calculations in order to permit resolution of such high frequency

variations.

5-3.2 Leeward Pressures.

BID predictions of engine face total pressures in the leeward

inlet are also generally in fairly good agreement with the test data

with respect to the maximum blast pressure level and the duration of

the principal pressure pulse (see Figures 5.11, 5.13, and 5.15). However,

it is evident that the BID pressures noticeably lag the experimental

pressures with a noticeably slower initial rise of the BID pressures

to their peak values. These two related differences may be attributed

to the fact that the BID code is a two-dimensional code which assumes

that the essentially side-on blast wave for these runs can reach the

leeward inlet only by the relatively diffuse process of diffraction

around the apex point of the inlet ramp (see Figure 1.1), whereas in

actuality the blast wave can also enter the leeward inlet more rapidly

by passing directly over or below the outboard inlet. In addition it

should be noted that the blast input characteristics used for the BID

calculations (Figures 5.1 to 5.3) may not be as applicable to leeward

See footnote on preceeding page.

142

inlet calculations as for blastward inlet calculations, since the

incident blast wave could be distorted significantly by fuselage inter-

ference effects before reaching the inboard inlet.

5-4 DISTORTION AT ENGINE FACE.

Engine face distortion time histories computed from the BID

code runs are compared with the corresponding AEDC test data in Figures 5.16

to 5.21 for the same runs discussed above. Test data and BID calculations

are shown on left and right hand sides of facing pages, respectively.

Since the BID code is a two dimensional code and the distortion

definitions normally used in B-1 studies (see Reference 1) are three

dimensional concepts involving pressures at 40 locations at the engine

face location, it was necessary to relate the BID cell pressures to the

pressures at the 40 locations. This was done for BID computations

simply by taking the BID pressure at each of the 40 engine face locations

to be equal to the pressure in the BID cell within which the corresponding

engine face transducers is located.

In Figures 5.16 to 5.21 the time of blast arrival at the blast-

ward inlet is indicated approximately by the start of the BID curves on

the right hand pages and the time of blast arrival at the engine face is

indicated by the start of the first ripples or ramplike rises of the

distortion parameters.

5-4.1 IDC.

Consider first the circumferential distortion parameters IDC1

through IDC5 for the individual engine face rings 1 to 5 in parts "a"

of Figures 5.16 to 5.21. The pre-blast values of the BID distortion

coefficients (before blast arrival at the face) are seen to be qualitatively

in good agreement with the corresponding test data for both blastward

and leeward inlets to the extent that the distortion increases signi-

ficantly in going from the inner ring (IDCl) to the outer ring (IDC5).

Quantitative agreement is only fair, with the experimental distortions

being generally larger than the BID computed distortions.

See Reference 1 for an explanation of distortion parameters and otherterminology used here.

143

After blast arrival both theory and experiment indicate

circumferential distortion increases generally lasting at least several

milliseconds. The details of distortion behavior and correlation are

somewhat different for the two inlets and for the different runs as

discussed below.

For the blastward inlet for all three runs (Figures 5.16a,

5.18a and 5.20a) the computed blast-induced circumferential distortions

are seen to be somewhat similar to the experimental distortions for the

outer engine rings (IDC4 and IDC5) but the experimental distortions are

larger than the computed distortions for the inner rings (IDCl and

IDC2). These larger experimental distortions can probably be attributed

partly to the basically three-dimensional effects of the bullet nose

hub of the model engine which is not taken into account in the two-

dimensional BID code.

For the leeward inlet for Run 39 there is somewhat better

correlation between theory and experiment for the inner engine rings

(Figure 5.17a). Here, after blast arrival, both calculated and experi-

mental distortions tend to rise to similar Levels for the inner rings

(IDCl and IDC2). However, the calculated distortions are too large for

the outer rings (IDC4 and IDC5).

For the leeward inlet for Run 40 about the same trends are

observed as mentioned above for Run 39 (see Figure 5.19a) except that

all the computed BID distortions rise to clearly too large values at

late times. This large predicted distortion might be attributed to the

fact that the BID code does not contain an adequate simulation of

viscosity effects for the circumstances of this run.

For the leeward inlet for Run 18 (Part 624) there appears to

be fairly good agreement between the experimental and calculated dis-

tortion trends for the relatively short time of the calculations.

Calculated and experimental time histories of the total cir-

cumferential distortion parameter IDC may be compared in parts "c" of

Figures 5.16 to 5.21. This parameter is a weighted combination of the

144

individual values of IDCl through IDC5 discussed above. Generally

the calculated pre-blast values of IDC are too low. The calculated

blast-induced values are too low for the blastward inlet for Runs 39

and 40 but are too high for the leeward inlet at late times for these

runs. For Run 18 calculated and experimental IDC values are similar

for both inlets.

5-4.2 IDR.

The engine face radial distortion parameters IDRl through

IDR5 and the total radial distortion parameter IDR are shown in parts

"b" and "c", respectively, of Figures 5.16 to 5.21. Generally the

experimental distortions are seen to appreciably exceed the BID-computed

distortions, particularly for the outer ring (IDC5). Some of the dis-

tortion variations are qualitatively similar, e.g., the rising distortion

variations for IDR4 and IDR5 in Figure 5.20b, but for other cases, e.g.,

Figures 5.17b and 5.19b, the experimental distortions indicate sub-

stantial blast induced distortion drops which are not predicted by

the BID code.

The differences observed here may be attributed primarily to

three-dimensional effects and viscosity effects not covered by the

BID code.

5-4.3 IDL.

The overall distortion parameter IDL is a weighted combination

of the TDC and IDR parameters (see Reference 1). The calculated pre-

blast values are generally substantially less than the experimental

values (see parts "c" of Figures 5.16 to 5.21). The calculated blast-

induced values are generally lower for the blastward inlet (Figures 5.16c,

5.18c and 5.20c) and they appear to become too high for late times for

the leeward inlet (Figures 5.17c and 5.19c).

5-4.4 IDA.

The calculated pre-blast average distortion parameter IDA is

generallv substantially less than the experimental value (see parts "c"

of Figures 5.16 to 5.21). The calculated blast-induced distortions

145

are also lower for the blastward inlet (Figures 5.16c, 5.18c and 5.20c)

but reach levels similar to the experimental values for the leeward

inlet (Figures 5.17c, 5.19c and 5.21c).

5-4.5 IDT.

The calculated pre-blast total distortion parameter IDT is

generally substantially less than the experimental value (see part "c"

of Figures 5.16 to 5.21). The calculated blast-induced distortions are

also less for most conditions except for the inboard inlet at late times

for Runs 39 and 40 (Figures 5.17c and 5.19c) where the late-time cal-

culated and experimental distortions are similar.

5-5 CORRELATION SUMMARY.

In summary, it may be concluded from the preceeding comparisons

that the BID code provides a good representation of those features of

the blast-induced inlet flow which can be reasonably represented by a

two-dimensional inviscid approach, particularly the inlet ramp and cowl

pressures and engine face pressures on the blastward inlet. For flow

characteristics which may be appreciably affected by three-dimensional

effects, such as the leeward inlet and engine face pressures, agreement

is not as good but still fair. For flow distortion characteristics,

which may be affected both by three-dimensional effects and by viscosity

effects (not included in BID), the agreement is less satisfactory. The

BID code generally underpredicts the pre-blast experimental distortions.

For the blast induced distortions, the code generally substantially

underpredicts distortions for the blastward inlet. For the leeward

inle', the code underpredicts IDR, and, for some cases, predicts too

high values of IDC and IDL at late times.

The comparisons would be expected to be generally improved by

extension of the BID code to three dimensions, particularly in regard

to distortion. This extension would require a significant increase in

the computer storage and computation time requitements. The rapid

strides in improvement of computer capacity and speed makes this

extension to three dimensions feasible now.

146

8

6V

0. 001

a) 44

950 -

44 850L

-H 40

40> CI

0

-14

6

w CO 4Q) -

*0 C/ 0

950

850 i

400

0 4 8 12 16 20

TIME (MSEC)

Figure 5.2. Blast input characteristics for Run 40 (Part 619).

148

CD____/ _

~7)

$.4 6oJ

~4

ht

0.004., O.01r

'- .. '

8 0-

Cd

0 .

" L,0 I I I

8 00 r

CU11 700 F L . ~ .

o 4~oO

0 4 8 12 16 20

TIME (MSEC)

Figure 5.3. Blast input characteristics for Run 18 (Part 624).

149

1 .25S _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

a)

0.95L) 0.80

1.1

ca 0.95

2 0.65

- 0.95

1.2S

- 0.950.80

4 8 12 16 20TIME (MSECS)

Figure 5.4. Comparison of theoretical and experimental time historiesof ramp pressures for Run 39 (Part 544).

150

1.Z0__

1:2 -____

O.95

1.25

1.2s

LA 0.s

- 0.80

U. S0 11 .25 _ _ _ _ _ _ _ _ _

1. 1 a

m 0.80L

1 . 2S D _ _ _ _ _ _ __ _ _ _ _ _

1.151

1.25BID11d

1.95C'dtTLst:Dt

1.21

- 1.95 L0I .81L

1.25

1.10

1.801.65

1.51

- 1.15

- 1.95

0)13

6 12 16 21TIME CMSECS)

Figure 5.6. Comparison of theoretical and experimental time historiesof ramp pressures for Run 40 (Part 619).

152

1.*25

N 1.951.80

- 1.65

1.11

0.95

- 1.25

1.1

3.95

1.21

1.11

1.95

8 .65

1 .25

N 0.95

'46 12 is 21TlIME CMSECS)

Figure 5.7. Comparison of theoretical and experimental time histories

of cowl pressures for Run 40 (Part 619).

1.'41

1.25

L 1.2S

Uv)

(0.d'% - Cd

O.I) I

- 1.19

0.95

00

1.2S4

- 1.13

3.154

1 .25

1.1.

Cv 1.95

0.80

1. 10TetDa

Cvn 0.95

1.25S _ _ _ _ _ _ _ __ _ _ _ _ _ _ _ _

0.95

8 .65

1.50

1.10

0.95

0.65

1.10

N1 0.95

N 0.65

O . SI) _ _ _ _ _ _ _ _ __ _ _ _ _ _ _ _ _

'4 8 12 16 2nTIME (MSECS)

Figure 5.9. Comparison of theoretical and experimental time historiesof cowl pressures for Run 18 (Part 624).

155

1.11

BID Code'~-Test Data

1.33

1.23

1.1

1.33,

I 1 12 114 lie1 23TIME (MSECS)

Figure 5.l0a. Transducer 1807.

Figure 5.10. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 39 (Part 544),blastward (outboard) inle2t.

156

1.413

BID Code

1331 A-~~ Test Data

1.93

@.Sol6 S 3 12 14 S 1 23

TIME CUSECS)

Figure 5.10b. Transducer 1810.

157

1.33 BID Code''--Test Data

1.93

C12 1.12

TIME (MSECS)

Figure 5.10c. Transducer 1812.

158

l.7-1

BID Code,-' Test Data

'.159

BID Code

1 33 -~-~--~--Test Data

1.23

OD

3.93

3.A3N6 3 12 14 Is is 23

TIME (MSECS)

Figure 5.10e. Transducer 1827.

160

1.1

BID CodeSTest Data

1.161

BID Code

1.33'-'-'--Test Data

1.262

BID Code*--- Test Data

1.21

1.1

1.33,

6 3 12 14 is Is 23TIME CMSECS)

Figure 5.10h. Transducer 1835.

163

BID Code

1.33 1~-' Test Data

1.2

6 312 1'# 1is 23TIME (MSECS)

Figure 5.11a. Transducer 2807.

Figure 5.11. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 39 (Part 544),leeward (inboard) inlet.

164

BID CodeivA-''- Test Data

1.21

1.33

6 IN1 12 111 is 18 21TIME CMSECS)


165

BID Code

~--Test Data

1.21

1. 13

1.336

BID Code1.33 '~- Test Data

CD

Figure 5.11d. Transducer 2815.

167

BID CodeTest Data

Cq

TIME (MSECS)

Figure 5.lle. Transducer 2827.

168

1.38 BID Code---- Test Data

O. D

6 0 12 1'4 16 18 20TIME CMSECS)

Figure 5.11f. Transducer 2830.

169

- BID Code

1.38~ Test Data

3.9,

8.88D10 i 12 14 is 1 23

TIME (MSECS)

Figure 5.11g. Transducer 2832.

170

BID Code

f'---Test Data

1.20

U).1.

lGD

U. 8

6 8 10 12 14 16s 1 20TIME (MSECSJ


171

BID Code1.33 Test Data

1814is is3

1.172

BID Code

1 .38 ~" Test Data

3.90

*TIME CMSECS)


173

BID Code

' ' Test Data

1.31

CO

68 132I 14 is 18 23TIME (MSECS)


174

BID Code

' - Test Data

1.33

1.2

3.93

3.83 1is 1 12 14 is is 23

TIME (MSECS)


175

BID Code

STest Data

1.31

1.190OD

1.93

6 8 13 12 1'I 1s 18 20TIME (MSECS)


176

BID CodeTest Data

1.23

1.13 1

1.177

1.141

BID Code~--- Test Data

1.33

6 8 3 12 14 16 1 21TIME (MSECS)


178

1.41

1.33 BID CodeTest Data

in

1 1.13 2

TIME (MSECS)


179

1.413

BID Code

STest Data

1.21

1.13

6811 12 16i 113 23TIME (MSECS)'



180

1.413-- BID Code

'-Test Data

1.33

CD

1.33

ILI

S S 1 12 16 is 8i 21TIME (MSECS)


181

BID Code'- Test Data

Cq

1.33

TIME (MSECS)


182

1.13

BID Code

1.33 "'-~'-Test Data

1.O

3.183

BID Code' Test Data

1.31

1.21

( 1.13GoCq

1.91

1.61 I -6 1 1I 12 1' 16 l6 21

TIME (MSECS)


184

BID Code'-'"--Test Data

I1.3e3

1.213

c) 1.1,

1.383

3 6 8 13 12 16 Is i 21

TIME CMSECS)


185

BID CodeTest Data.

1.33.

0

Cq

TIME CMSECS)

C Figure 5.13g. Transducer 2832.

186

BID Code133 ~Test Data

CO

6 3 12 14 is is 23TIME CMSECSJ


187

BID Codelest Data

OD S

IL

0.888

AOAOBO 511 KAMAN AVIDTNE BURLINGTON MA F/B 21/5F URTHER EVALUATION OF BLAST TESTS OF AN ENGINE INLET U)

UA 79 J UTNK LT ATNYO DAO 8C03UNCLASSIFIED KA-TR-165 DNA-4994F NL

-mEhE3E-EhhEhEinnunnunuunnuu-mEEhhEEEE

-- BID Code~-Test Data

12 4 640 2

Fiur 5.4.TrndceU30

189

BID Code*-Test Data

19

BID Code-.- Test Data

1.40D

W" 1.0

6 8 10 12 1'4 118 20TIME CMSECS)


191

-- BID Code4$v-Test Data

N 1.20

1.0

I . 6oe6 is le 12 1'4 6 18 20

TIME (MSECS)


192

BID Codei w Test Data

roe 6

6 8 18 1 1'4 6 18 20TIME (MStCS)


193

BID CodeSTest Data

0

6 82 ID1Z 4 1 is1 23TIME CMSECS)


194

BID CodeNW"-Test Data

1. 3D

TIME (MSECS3


195

BID Code

1 .6 4-,-- Test Data

N*

1.20

-- 1. Ut

I.II I

6 6 13 12 14 16 16 20TIME (MSECS)



196

1 * GDBID Code-'- Test Data

CD

IAI

S 0 12 1'4 16 Is 20TIME (MSECS)


197

____BID Code

1 * 61 - Test Data

0

6 S ID ia 1'4 is 1S ZTIME (MSECS)


198

BID Code

I 1 GO' Test Data

O. S

S S 10 1ZI4 is I 20TIME (MSECS)


199

____BID Code

1 * 0 -w--~---.Test Data

032

048

6 8 tO 12 1'4 6 18 26TIME CMSECS)


200

BID Code1 . 0 t-~-~Test Data

1.40

0.601

BID CodeI. GO0~-v~ Test Data

Ct)

6 312 14 is is 20TIME (MSECS)


202

BID Code

* .~-w'~Test Data

0

ID 12 1'4 6 S 20TIME (MSECS)


203

OUTBORRO INLET

0.25

0.20

0.15

0.05 _,__ __,h_. ___ I1 __,. _

0.00- ____

0.25

0.20

c* 0.15

en 0.10 -_

0.05 f, I . ,, 1 , -

0.000.25

0.20

Mn 0.15

- 0.05 NJJAW #

0.000.25

0.20

L 0.15 -

en 0.10 -

- 0.

0.000.25

0.20 --

uo 0.15I

0.00'0.00 4.00 8.00 12.00 16.00 20.00


EXPERIMENTAL DATA

Figure 5.16a. Circumferential distortion.

Figure 5.16. Comparison of theoretical and experimental time historiesof engine face distortion for Run 39 (Part 544),blastward (outboard) inlet.

204

.25

.20

.15

(aso .1 0

.25

.20

N' .15

Uo .10

.25

.21

en .15

-4 o. Oo S

.25

.20

.15UC3

0.00

U. D

.25 -_ _ _ _ _ _ __ _ _ _ _ _ _ _

.25

.20

_ .15_

U .10

S .05

0.000 I I I

0 4 8 12 1

TIMECMSEC)

BID CODE

Figure 5.16a. Concluded.

205

OUTBORRD INLET0.100.08

- 0.06

0.04

0.02 I "

0.00

0.08

CJ 0.06

C2 0.04

0.02 I

0.00 *,Li0.100.08

mr 0.06

CD 0.04

0.02

0.00

0.06 .. . .

C2 0.04

0.000.10

0.08Lf) 0.06

23 0.04


EXPERIMENTAL DATA

Figure 5.16b. Radial distortion.

206

.02

10

.02

.10

.02

.08k.08k

.02k

o ~ .04

.02

0.00 4 a 12 16 21

TI ME(MSEC)

BID CODE

Figure 5.16b. Concluded.

207

OUTBORRD INLET

1 .50

0.60 nj, ", .1 P.

0.30 ., - , - - -

0.000.25

0.20------..j-

0.15Xi IL_- --

0.000.10

0.08 - - - - _ -

OI rk

0 .02 .. .. .. .. .. . . i ..

0.001 .00

0 .8 0 . . ... . . .. . +- .. . .. . .

r-0.60 - ~ --- --- ~ -

CD 0.40 ' . . . . . . .

0.20 ... .

0.00

1.000.80 --- _ _ _- _ __.

0.60

C:' 0.40

0.000.00 4.00 8.00 12.00 16.00 20.00

TIME IN MILLISECONDSEXPERIMENTAL DATA

Figure 5.16c. Overall distortion.

208

1.20

. 90

.30

.20

0 * I.0

as8

Q 20

.80

.20

.40

.20

0.0000 '4 8 12 16 20

TIME CMSEC)

BID CODE

Figure 5.16c. Concluded.

209

INBOARD INLET

0.25

0.20 10.15 -

C] 0.10 _ __ _I'

0.05 A

0.000.25

0.20

Nj 0.15 t

0.05 -V

0.00 _ _ _ _ _ _ I "' i'li m0.25

0.20 --Mr 0.15

-10

0.05 -1.----v-0.000.25

0.20 __-~ -- ~- _______

"* 0.15 - _

D- 0.10 1 70.05

0.000.25

0.20 { -

Lro 0.15

C: 0.10

0 .05

0.000.00 4.00 8.00 12.00 16.00 20.00


EXPERIMENTAL DATA


Figure 5.17. Comparison of theoretical and experimental time historiesof engine face distortion for Run 39 (Part 544), leeward(inboard) inlet.

210

rJ

.25 ------

*20

atsa

.25__ _ _ _ _ _ _ _

i s

0.110

.25 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

.2a~cr) IF)

02

us

0

.2s

. 20

Cr)

a-- 12 1620

TIMECMSEC3

BID CODE


211

INBORRO INLET

0.0

•- 0.06.04

___o_____ --._ ___-V __0.000.10

0.08

cJ 0.06 - _-_

r- 0.04

0.02

0.000.10

0.08 -- -- _

S0.04f I

0.000.10

0.08 -- __ __ _

0 .06 .. ... . .. . . . . .

C 0.04 .

-0.02 ~.A

0.10

0.1o8

0.08 Li

C030.04 (jJ "i~ii~ A~'0.02

0.000.00 4.00 8.00 12.00 16.00 20.00


EXPERIMENTAL DATA


212

0.0I

.10

.08

.02k_, .06

cr_

.02 L

.n2

.08 _

C 08

r_

04 _

.02 _

. 0 12

.08k

.06_

o7 .]4 1_

F~. .102 __ _ __

.08

:U- .06 _

c .04

"" .02

0.00 I " I I,

0 4 8 12 16 20TIME(MSEC)

BID CODE


213

" k i I I2

INBORRO INLET

1 .20 ____1l.~

0.90_: _ 0 .6 0 1

0.30 __I I0.000.25

0.20

0.152 o.'o T- _ ... _ "

0.05 A, WA0.00

0.10,o0.08

0 .0 4 . . . . .. .1I . .. .. ..

0.02 ...

0.001 .00

0.80 ___

0.60 - _ 4,C30.40- _ _

0.20

0.00

1 .00

0.80 _ - _

0.60 ...

o 0.40

0.20 - _ - - - _

0.00 I I0.00 4.00 8.00 12.00 16.00 20-00


EXPERIMENTAL DATA


214

1. 2 0 L -_ _ _ _ _ _

-J.90

.25

. 20

.10

o ~ .Lf

.02

. 80

- .20

0.00 L

1.00 -_ _ _ _ __ _ _ _ _ _

. 80

.Go0

C2 40l

.20

0 .0 00 '4 8 12 1 G 20

TI MECMSEC)

BID CODE


215

OUTBORRO INLET

0.25

0.20

--' 0.15

2 0.10 ?0.00 ki

____

0.25

0.20

0.15

- 0.05 VO

.00

0.00 o* , , V .,. , _____________.....__

0.25

0.20

') 0.15

-3 0.10 ____ ______0.150.000.25

0.20

0.15

2D 0.100.05 A A Ititrf' , rw0.00

0.20

u' 0.15

0.000.00 4.00 8.00 12.00 16.00 20.00


EXPERIMENTAL DATA


Figure 5.18. Comparison of theoretical and experimental time historiesof engine face distortion for Run 40 (Part 619),blastward (outboard) inlet.

216

.28

.23

- .15

P . go

.21

o .10- .15

.25

.21

.15

o .13

.25

.28

.15

0.10

.20

LP .15U-0

3.333 4 612 16 2

TIMECMSEC)BID CODE

Figure 5.1.8a. Concl1uded.

217

OUTBOARD INLET

0.10

0.08

- 0.06

0.04

0.02 _,JILL ii0.000.10

0.08

CI 0.06

C3 0.04

0.02

0.00

0.10

0.08

' 0.06

0.04

0.020.00 I .... m... .

0.10

0.08

, 0.06o: 0.04 f .

0.02 r Til rjT--~ ~ . 4 '+'~iit .j

0.000.10

0.08{

0.06 I 11

- 0.04

0.00



218

.16

.06_

0: ....

o o

.82

.10

DDu t ,I I I

* D

.08.84

.02

- . .2

uOu DDIt I

" .02

uTuI I EI MSC.18

.18.0D2

D.D Be*

a1 __ __ _ _12__ __ _ __is__ __ __ _21___ _

TIMCusC

ID CODSEC


219

OUTBORRO INLET1.50

1.20

0.90 -. __ __ _

C3 0.60 ___1 WZ Aki4 tJ0.30 f____0.000.25

0.20

0.15

0.00 ____

0.10____ _

0.08 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

0.06

0.04 Ihhdj r0.02-[

0.00 -

1.00-

0.80

0.60

2: 0.40

0.20

0 .00

0.80

0.20

0.00 _____


EXPERIMENTAL DATA


220

1.23

. 93

.30

.21

o .10

0.10

.138

.12

.63

- .20

1.63

.83

.23

3.033 4 IN 12 16 21

TIMECMSEC)

BID CODE


221

INBO:RO INLET

0.25

0.20

0.15

2 010 -L] 4

0 .00 "" . . . • *

0.25

0.20

0.15 i i j).L , w ___,____2l 0.10

0.05

0 .00 - • "_ -0.25

0.20 --

') 0.15

00 _____ ,____ _ ______ _____

0.25

0.20 - _ _ _ _ _ _ _ _ _ _ _

0.15 Jul,

0.000.25

0.20

LI) 0.15

0.0120 000 000.,o .o,.o o oo

TIME IN MILLISECONDSEXPERIMENqTAL DATAFigure 5.19a. Circumferential distortion.

Figue 519.Comparison of theoretical and experimental time historiesof engine face distortion for Run 40 (Part 619),leeward (inboard) inlet.

222

.25

.21

2.18

L)o .11

- . go

.2S

.1

U .10

- . Be

3.2S

.21

.15

Po .15

.2S

.21

C- Io3 al

AuSL

1 4 1 12 to 23

TIMECMSEC)

BID CODE


223

0.1 _____INBORO INLET

0.08

0.06D0.04

0.00

0.020.10

0.08

CJ 0.06C3 0.04

0.02-

0.00 lJ AM Li0.10

0.08

m0.06o0.04

0.02

0.00.10

0.08

** 0.06

S0.04 ____1

0.0 li t____ "l________

Ll 0.060.10__ I

o0.04 8iA

LO0.003 0.00 4.00 8.0 20 6000

0-224

o .s

.13

.36

.04

1.13 0

o B

.3

3.30

.36

- .12

LO .46 21,2

TI ME CMSEC)

BID CODE


225

INBOARD INLET

!.50

1.20

0.90 "" ' j'

0.00

0.25

0.20

0.15

2, 0.10

0.05

0.000.10

0.08

0.06

2 0.04

0.021A

0.001.00

0.80

0.80

0.60

0.20

0.000.00 4.00 8.00 12.00 16.00 20.00



226

1.23

.90

C33

.25 __ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

.20

p.-5

.13

ct* 06

o- .02

S. of

.91

- .21

O * 12t32

.227

OUTBORRO INLET

0.25

0.20

- 0.15_ _ _ I i2 0.10 )IT r !r !S

0.05 1:, i$[I } il I

0.00 _____rI

N0.25

0.20-

2R 0.10

0.205

M 0.15 I-20.10

0.00L__ _ _ _

0.25T

0.20 _ _ _ 1_ _ _ _ _ _

LI) 0.152: 0.101 W

0.05 20.00

0.00 402.0150 60 00TIEINMLLSCOD

EXEIMNA DATAFiue52aUicmeeta itrin

Figure gur 5.2.2omario ofrctheorential adiexperienta.iehitre

of engine face distortion for Run 18 (Part 624),blastvard (outboard) inlet.

228

.25

.20

.2S

.20 I a

co .15

_.LU _

O.* 25 _ _ _ __ _ _ _ __ _ _ ____ _ _ _ _ _ _ _ _ _

,25

.20k

C.1,

o .10

0 .005 _ _

.2U

0.20_ 0.00 I I I I

14 8 12 1620

TI ME MSEC)

BID CODE

Figure 5,20a, Concluded.

229

OUTBORRO INLET

0.10

0.08

-0.06

C20.04 ]A0.02

lt

0.040~0.10

0.08

('j 0.04

0.08

S0.060

20.10 if ~ -

0.02

m0.00

0.04

0.02 _______i. iI0.00

0.00 .0 80 20 60 00

0.230

S04

.08L

CI I,131

.02

0.00 I I I I

.08 _8

.04

.02

08 _

.02_

0].O0 L._.-.- .

U G.02 C

.I] _41

T I ME C MSE CBID CODE


231

OUTBORO INLET

1.50

2 0.60141; I_ _

C30.30

0.000.251

0.20

0.15 4

0.0504 4 ~ I~1___0.10

0.08 - _ _ _ _ _ _ _ _

0.06

C, 0.0C4 iil1

0.02

0.001 .00 _____

0.80

CC0.60

0.00____ _

1.00__ _ _ _

0.80

0.60

CR 0.40__ _ -- if0.00 _____


EXPERIMENTAL DATA


232

1. Inv._ _ _ _

1.20k

.go

.30L

.25

.20f-

.Bs

,08

*06

.02

Ua a _ _

.810

.GD

<]i0

.20

a.80 a

.00 8 1212020

TIMECMSEC)

BID CODE


233

0.25

0.20

- 0.15

2C3 0.100.0o .os I, ,,..i ll), . , ., ,. . .,,,

0.25

0.20

C\I0.15

2 0.10 .,

0.000.25

0.20

Cr 0.15(..)

2 0.10

0.05 j ''iF.t' I 1 *.O"

0.00 OW

0.25

0.20

V, 0.15 Jj

62 0.10 -t0.05 'ti.t '

.0.000.25

0.20

LID 0.15

0.05 4v 1u rv * r

0.000.00 4.00 8.00 12.00 16.00 20-00



Figure 5.21. Comparison of theoretical and experimental time histories

of engine face distortion for Run 18 (Part 624),leeward (inboard) inlet.

234

.25

.20

.2o L

n_ .l a

.25

.20V

uLF

0 .D L I

.25 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

Figur 5.1. ocldd

* 230

I-. .

*, . L * J -- '-I-

.1LV

i 0 0 .L 4

U 812 1G 2

BID CODE


235

INBORRO INLET

0.10

0.08

O.0 'IC 0.04

0.02

0.10-

0.08

(Y) 0.06

c 0.04

0.02 -_"_

0.10

0.08

*r) 0.06

o 0.04

0.02 -4 lta

0.00 ..... LJ . . ..

0.10

0.08

0.06

0.00

0.02

0.000.10

0.08

0202

0.00 4.00 8.00 12.O0 16.00 20.0WTI!ME I N MI]LLIS ECONDS

EXPERIMENTAL DATA

- Figure 5.21b. Radial distortion.

: 236

* oe.10

o02

04O

2-0

LL .°04 f

00

0.00 I I i

. 02

c-- .F02

. 08 L

_7 .04- -02.02

0.00 I I ____

U4 8 12 lG 20TI ME CMSEU J

BI:D CODE


237

INBORRO INLET

1 .50

1.20

0.90

23 0.60 ~I

0.30 '_____

_____

0.000.25

0.20

0.15-___ ____

____

0 0.10 -____ ________

0.05 '

0.10

0.080.06 1jJ

0.02

0.80 _ _ _

C: 0.40

0.20N-6

0.00____ -

0.0040080120160200

0.238

1.20

.90

.30

. 25 __ _ _ _ _ _ _ _ _ _ _ _ _ _ _

. 20

. 15Uo .10

0.0 as L

.08

0 .04

.02

.80

os

- .20k

a 0

08 12 is20

TIMECMSEC)

BID CODE

Figure 5.21k. Concluded.L - 2394

SECTION 6

EVALbATION OF LATE TIME LARGE DISTORTION VALUES

It was found in the inlet-blast tests of Reference 1 that

generally very few large distortion values were obtained during the

early-time, definitely blast-type, flow periods of the tests. However,

there were some observed large distortion values at times after the

fairly definite blast type flow duration of about 3.3 ms, which appeared

to deserve more detailed study.

As part of the present study, a detailed re-evaluation was

made of the 16T inlet blast test results to determine whether the large

distortion values observed at late times (after the nominal blast event)

on some firings could be attributed to inlet response behavior or to

other effects. An initial appraisal identified 26 firings with signi-

ficant late time distortions which appeared worthy of consideration.

For these runs it was found that the large observed IDL values could be

correlated with one or more of the following circumstances.

1. Some IDL peaks correlated closely to significant rises in

total pressure, indicating that these peaks were produced

by the effects of the cold jet from the shock tube

passing directly into the inlet.

2. Some IDL peaks corresponded closely to rapid changes in

input pressures measured by the claw probes, indicating

that the distortion values were caused primarily by input

variations rather than inlet response effects.

3. For some runs, there were significant differences in the

input pressures measured by the different claw and static

probes, indicating a considerable distortion of the flow

entering the inlet. Hence, for these runs the engine

face distortion could have resulted from the input

distortion rather than from inlet response effects.

240

4. For several runs, large apparent distortion peaks were

caused by the readings of one pressure transducer (1802),

whose readings were somewhat erratic and were inconsistent

with the measurements of two adjacent transducers (1801

and 1804).

5. For at least one run, large apparent distortion values

were attributed to erratic behavior of a group of 4

pressure transducers which were recorded on the same

oscillograph track. These transducers gave pressures

inconsistent with adjacent transducers.

6. In some cases transducers bottomed, producing false

indications of large distortion values.

In summary, no cases were found where large late time distortion

values (IDL > 1) were obtained which could be definitely attributed to

inlet response behavior. In all cases there was evidence to suggest

that the large distortion values could be attributed at least in part to

either cold gas jet impact on the inlet, large input distortions or

transducer malfunctions.

The above observations do not completely resolve the question

as to whether any of the large late time distortions observed in the

tests can be attributed to inlet blast response effects. It can only be

said that the factors pointed out above would appear to make any such

determination from the late time test data a difficult matter and any

results of such a determination could be suspect.

241

-I. .' ' . . - 7 . . - - : '- ... . ........ > ' ...,, .. . : -iF .... -- _

SECTION 7

REFLECTED SHOCK-BOUNDARY LAYER INTERACTION

The reflected wave from the engine, as it moves upstream

through the inlet, makes an adverse piessure gradient for the flow

within the inlet. The adverse pressure gradient produces some dis-

tortion of the flow. If the gradient is large enough, boundary-layer

separation takes place and large distortion of the flow would result, as

pointed out in Section 3 and 4 and in Reference 1.

A sketch illustrating the effect of the reflected waves on

the boundary layers in the two inlets is shown in Figure 7.1. The waves,

shown here as two shocks, produce an adverse pressure gradient within

the boundary layers on the walls of the inlets. The pressure gradients

produce more rapid thickening of the boundary layers and retardation of

the flow within the layers. If the gradients are sufficiently large,

separation of the flow from the walls would take place resulting in

large distortion.

There were indications that separation by the reflected shock

may have taken place in some of these tests. The problem is that distortion

at the engine face would have resulted after the end of the test period,

about 3 ms, so effect of separation could not be verified.

Calculations are presented in Reference 1 using the NASA

BLAYER code of Reference 4. As pointed out there, the BLAYER calculations

are limited by the assumption of a constant total pressure and total

temperature in the free stream, whereas the variations produced by blast

interaction can be large (Reference 1).

NASA has generalized the BLAYER code, subsequent to the work

published in Reference 4, to include variations of the total pressure

and total temperature in the free stream. A copy of the generalized

BLAYER code was provided to Kaman AviDyne (Reference 5), and it has been

employed for the calculations reported below.

242

7-1 SHOCK-BOUNDARY LAYER CALCULATION.

Calculations of the boundary layers on the inlet cowls and

ramps were made using pressure, temperature and velocity data from BID

code results for the free-stream conditions. BID results for properties

in the cells contiguous to the respective wall were employed. The BID

data used were the same as for the calculations reported in Reference 1:

BID Run 10/26/77, M=0.85, W2R=350 lb/s, 4=90 deg and Aps=5 psi. The

data selected were for a fixed time of 33.6 ms (full scale) after blast

arrival, when the reflected waves were well into both inlets.

The generalized BLAYER code assumes a steady-state boundary

layer, i.e. the conditions do not vary with time. Therefore the BID

data distributions for the free-stream properties were assumed to be

frozen, i.e. non-varying with time. In actuality the reflected waves

move upstream with time, so the properties are time varying. The

assumption of a steady state flow is believed to be conservative, because

the pressure gradients experienced by the boundary layers with the

moving wave would be greater than for the assumed frozen conditions.

The results of the calculations are expressed in terms of a

parameter called the boundary-layer factor Hi, essentially representing

the velocity distribution across the boundary layer (normal to the

wall). If the velocity distribution is more uniform, the parameter

approaches unity. A value of 1.2 to 1.4 is typical for a turbulent

boundary layer on a flat plate in the absence of a pressure gradient.

In an adverse pressure gradient (increasing pressure in the direction of

flow) the velocity decreases towards the wall so Hi increases. When the

velocity gradient normal to the wall goes to zero, the flow can separate

from the wall. Experience shows that separation may occur for Hi

values as low as 2.0 and definitely by 2.8.

The generalized-BLAYER calculations were carried out for three

cases: the boundary layers on the two cowls and on the blastward

splitter ramp. The boundary layers were assumed to be turbulent from

the leading edges.

243

Calculations were not made for the leeward splitter ramp,

because of undefined starting conditions. The data indicate a possible

separation bubble near the leading edge with probable reattachment. The

boundary-layer conditions are not believed to be well enough defined at

this point for meaningful boundary-layer calculations.

The results of these calculations are presented in Figures 7.2

to 7.4. The input data (PI' Pt' Tt) and output (Hi) are presented as

functions of the station along each inlet, measured from the leading

edge of the splitter ramps.

Within the blastward inlet the primary part of the reflected

wave at 33.6 ms extends between ramp Stations 12 and 17 (ft), as indicated

by the rise in total pressure. On the blastward cowl there is an adverse

(positive) pressure (static) gradient between Stations 8 and about 13

due to recovery within the inlet. The pressure-gradient decreases

somewhat between about Stations 12 and 13.6, where the pressure gradient

of the reflected wave is picked up. The form factor H. increases to a1

peak value of 1.57 at station 11.7, which would be well below the

minimum separation value of 2.0. H. then decreases to 1.38 at Station 14.81

where the adverse pressure gradient of the reflected wave causes it to

increase rapidly, reaching the separation value of 2.8 at Station 16.8.

At the ramp the adverse pressure gradient begins at about

Station 10. H. is fairly constant to about Station 16 where the gradient1

increases markedly. The separation value of 2.8 is reached at Station 18.3.

In the leeward inlet the reflected wave has not had a signi-

ficant impact on the flow at this time (33.6 ms). There is some effect

over about the downstream half of the inlet but the effect is to reduce

somewhat the large negative gradient in total pressure produced by the

incident blast wave. The static pressure gradient produced by the blast

wave is large, so Hi rises all along the cowl, and the boundary-layer

separation value of 2.8 is reached by Station 20.6. The pressure

essentially levels off beyond that point, so the determination of

separation is not as definite as for the leeward cowl.

244

It is concluded that separation caused by the reflected waves

would have occurred in the blastward inlet and possibly in the leeward

inlet. There are several factors that are believed to make the cal-

culation conservative (under-predict separation). First, a fan stage is

expected to reflect a stronger wave than the choked throttle that was

employed in the BID calculation. Seccnd, the BID code is dissipative,

so adverse pressure gradients are expected to be somewhat under-predicted.

Third, the generalized BLAYER code applies for a steady-state boundary

layer, and the unsteady effect is expected to increase the tendency

toward separation.

Tests are needed having blast durations that are sufficiently

long to observe the separation and resultant distortion at the engine

face. The test period for these test conditions should be roughly

doubled.

245

BLAS, WAVI

SHOCK

SHOCK DISTORTION

1INL ET FAN

Figure 7.1. Boundary layer separation and distortion from fan reflected shock wave.

246

3. DI

-

2. t1t

1.50.....

O S I O 15 20 35STATION (FT)

T tlTt

t

l . to

3.63 Pto

0

S 13 is 2 2sSTATION (rT)

Figure 7.2 - Predicted free-stream properties and form factorH of boundary layer on cowl of blastvard inlet.Ater reflected wave enters inlet.

247

3.10

I.SI

I-t

0. 0

0.88

..... .. ....... ....... ........ I .. . .AAA'D AA.6

I I is252

STATION (FT)

Fiur .3 Prdi tedfe-tempoete n omfco

3.838

3.33

<-

Ixc_2.UD

a

U2.28 2SsT,^TioN crT)

t

t 0

1.61

3 1 -t3 232

STATION (rT)

Figure 7.4 -Predicted free-stream properties and form factorH of boundary layer on cowl of leeward inlet.Ater reflected wave enters inlet.

249

Iwill

SECTION 8

CONCLUSIONS

From evaluating the results of blast-wave engine-inlet inter-

action tests with a B-1 type engine inlet, the following conclusions are

reached.

1. The blast interaction with the windward inlet produced a rise

in the mean total pressure at the engine face followed by a

decay. A second rise occurred due to reflection from the

choked control vanes simulating the engine. The magnitude of

the first peak is found to be nearly 1.41 times as large

(+ll percent) as the increment in total pressure across the

incident blast shock. This ratio varies some with inlet

weight flow and Mach number but it is unaffected by changes

in shock overpressure and intercept angle, over the range

tested (76 to 110 deg).

2. The reflection of the blast wave from the engine is expected

to be a potential cause of distortion through boundary-layer

retardation. The rise rate approaching the second peak in

engine face total pressure increased with shock overpressure.

The magnitude of the rise to the second peak increased with

lower intercept angles (more head-on) and lower weight flows

(low engine thrust conditions) and was essentially unaffected

by Mach number.

3. The mean total pressure at the engine face rose more slowly in

the leeward inlet than the blastward inlet, following blast

intercept, and by a smaller amount. In many of the tests the

rise in total pressure was followed within a millisecond or so

by a rapid fall-off, attributed to possible separation of the

flow from the walls of the inlet.

250

4. The two-dimensional BID code provides a good representation of

those features of the blast-induced inlet flow which can be

reasonably represented by a two-dimensional inviscid approach,

particularly the inlet ramp and cowl pressures and engine face

pressures for the blastward inlet. For flow characteristics

which may be appreciably affected by three-dimensional effects,

such as the leeward inlet and engine face pressures, agreement

is not as good but still fair. For flow distortion characteristics,

which may be affected both by three-dimensional effects and by

viscosity effects (not included in BID), agreement is less

satisfactory. To improve this situation, extension of the BID

code to the three-dimensional case appears feasible.

5. A study of large apparent distortion values observed at late

times during the AEDC tests, after the limited test period of

about 3 ms, indicated no cases where these large distortion

values could be definitely attributed to inlet response behavior.

The limited test period was too short to permit observation of

possible large late time distortion effects not masked by

extraneous factors. It is recommended that the test duration

for future tests be increased, by a factor of two or more.

6. Calculations of the effect of the engine-reflected blast wave

on the boundary layers in the inlets, made using the generalized

NASA-Lewis BLAYER code, agree with the previous results

(DNA 4590F) that indicated boundary-layer separation would

occur on the cowl and splitter of the blastward inlet and

possibly on the cowl of the leeward inlet. The blast shock

overpressure was 5 psi and the intercept angle 90 degrees.

251

REFERENCES

1. Ruetenik, J.R., and Smiley, R.F., Wind-Tunnel Shock-Tube Simulationand Evaluation of Blast Effects on an Engine Inlet, Kaman AviDyneReport TR-147, DNA 4590F, 1978.

2. AEDC Reports, Data Tabulations, Plots and Tapes on the DNA B-1Blast Effects Test, AEDC Project P41T-D2A, Test TF-419, 1976-1977.

3. Thompson, J.H., Tomayko, M.A., and Ruetenik, J.R., BID Code forComputing Aircraft Engine-Inlet Response to Blast Disturbances,Kaman AviDyne. Unpublished.

4. McNally, W.D., FORTRAN Program for Calculating Compressible Laminarand Turbulent Boundary Layers in Arbitrary Pressure Gradients,NASA Technical Note D-5681, May 1970.

5. McNally, W.D., NASA Lewis Research Center, Correspondence toDr. Ruetenik, 28 April 1978.

252

L i___ _ _

APPENDIX

TRANSDUCER CALIBRATION EVALUATION

The experimental ramp and cowl pressures measured inside the

subject model inlet during the tests of References 1 and 2 are of

interest both for assessing blast-induced inlet loadings and for pro-

viding a basis for evaluation of the BID code for predicting inlet

pressures and velocities. However, in using these data for such

purposes it is important for the data user to appreciate that some of

these data appear unreliable to some extent because of experimental

difficulties experienced during the tests. This appendix points out

briefly the principal problems encountered and indicates the degree

of reliability of different parts of the data.

The primary problem experienced with the ramp/cowl pressure

transducers was an inability to calibrate the transducers accurately

during the test period due primarily to unanticipated constrictions

in some of the tubes connecting the transducers to the calibration

pressure source. This problem could not be resolved in the very limited

time that was available for the model tests. The resulting data as

presented in References 1 and 2, therefore, had to be reduced on the

basis of sometimes nominal or questionable calibration results.

To clarify this calibration problem, KA examined the test

data for all ramp/cowl transducers with the aim of identifying questionable

data and providing correction factors if possible. Data calibration

errors were identified by such means as comparing transducer pressures

for the same transducer for similar runs, and/or by comparing transducer

pressures for adjacent transducers which should have about the same

pressure on the average. E.g., transducers 1970 and 1903 face each

other at the same axial location inside the outboard inlet, and normally

have quite similar late-time blast-response pressure time histories,

both according to many test runs and to the BID code. Consequently,

since transducer 1970 had no apparent calibration problem, any strong

253

difference between the indicated pressures for these two transducers

can be reasonably interpreted as an indication that the calibration

factor used for transducer 1903 is unreliable.

Using comparisons of this type, Table A.1 was prepared, which

indicates the KA estimate of the degree of reliability of the calibration

factors used to reduce the data presented in References 1 and 2. In

this table, questionable data are identified as either H or L, depending

on whether the pressure values presented in References I and 2 appeared

conclusively to be too high (H) or too low (L). Data for some other

runs also appears questionable to a lesser extent (not indicated in

the table), but the evidence for such cases is less conclusive.

Some attempts were made to obtain correction factors for some

of the questionable data, but it appeared that additional information

(not available for this study) would be required from AEDC to effectively

accomplish this purpose.

254

TABLE A.1

RAMP/COWL TRANSDUCER EVALUATION

AEDC Ramp/Cowl Transducer NumberPartNo. 1903 1905 1935 1950 1970 1990 2902

512 X X H X X H H513 X X H X X H H517 X H H X X H H518 X H H X X H H519 X X H X X H H525 X X H X X H H526 X H H X X H H527 X H H X X H H544 X X L H545 X X L H546 X L H550 X X L H551 X X L H553 X X L H558 X X H559 X X H568 X X H569 X H570 X H573 X H574 X H582 L H583 L H584 L X X H589 H H590 H H591 H H596 L X X X H597 L X X H598 L X X X H600 L X X H601 L X H602 L X X X H607 X H608 X X H615 X H619 X X H620 X H621 x H624 L X x625 L626 L X

*Transducers 1902 and 1980 appeared generally reliable; transducer 1904

provided no useful data.

X Designates no useful data obtained.H Indicates pressure values in Reference 2 appear to be too high.L Indicates pressure values in Reference 2 Pppear to be too low.

255

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Boeing Wichita Co. Kaman Sciences Corp.ATTN: R. Syring ATTN: D. Sachs

Calspan Corp. Los Alamos Technical Associates, Inc.ATTN: M. Dunn ATTN: P. Hughes

ATTN: C. SparlingUniversity of Dayton

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I

258 :

6