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FURTHER SHOCK TUNNEL STUDIES 

OF SCRAAIJFT PHENOAKNA 

NASA CONTRACT NO. NAGN - 674 

by: R.G. Motrgan, A. PLuLee, 

N.A. MOW&, R.3. SMkm 

RESEARCtf REPORT NO. 70186 
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NASA CONTRACT NO. NAGW - 674. 

R.G Morgan 

A. Paul1 

N. A. Morris 

R.J.Stalker 
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CONTENTS 

NOTATION 

ABSTRACT 

LIST OF FIGURES 

INTRODUCTION 

SECTIONS : 

A. Silane Combustion limits 

B. Pitot Rake Measurements 

C. Source Flow Nozzles 

D. Boundarv Laver Heat Transfer 

E. Film Cooling 

F. Multi-Staged expansions 

G. Shock Induced Ignition 

TABLE OF TEST CONDITIONS 

REFERENCES 

ACKNOWLEDGEMENTS 

APPENDIX 

FIGURES 

PAGE 

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7 

17 

21 

24 

26 

34 

44 

51 

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NOTAT ION 

C Specific heat at constant pressure P 

Stagnation enthalpy HS 

I Injector location 

I' Net specific impulse SP 

Bessel function of order v JV 

K Thermal conductivity 

L . Length scale parameter 

Intake Mach number 

Hydrogen jet Mach number 

MI 

MJ 

el Mass flow rate of injected gas 

P Local static pressure 

0 

Intake static pressure 

Stagnation pressure pS 

'IO 

Re 

ST Start of diverging section 

4 Surface heat transfer rate 

Flux of heat across fuel/air interface 

Freestream unit Reynolds number 

Intake static temperature 

Wall surface temperature 

Air velocity at intake 

Velocity of injected layer 

Wetted length measured from intake 

TI 

TW 

uI 

uO 

X 

Thickness of injected layer 

Y Cross stream co-ordinate 

Nth zero of Jv-1 

Thrust surface divergence angle 

'n 

OD 

-1 -1 

J kg K 

MJ/kg 

sec 

Wm -1 k -1 

m 

-1 -1 

kg s m 

kPa 

kP a 

MPa 

W cm-2 

Wm 

-2 

m- 1 

K 

K 

ms-1 

-1 

ms 

mm 

m 

degrees 
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V 

Equivalence ratio 

Insulating layer mass flux power law coefficient 

1 

3.+2 

P Density 

Density 

a Proport 

Po . of injected 

on of fuel 

r The gamma function 

layer 

n insulating .ayer 

AT Temperature difference between injected layer 

- and room temperature 

kg rn-3 

kg ~n-~ 

K 
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ABSTRACT 

This report is Dart of a continuing studv Qf scramjet related 

phenomena being undertakm bv the University of Oueensland. using the 

shock tunnel T3 at the Australian Naticinal Universitv. Simple two 

dimensional models were used with a combination of wall and central 

injectors. 

Silane as an additive to kvdrogen fuel was studied over a range of 

temperatures and pressures to evaluate its' effect as an ignition aid. 

At all these conditions the effect of silane concentration was 

investigated. Silane at concentrations sometimes as low as 2.5% was 

found til be effective in inducing ignition at conditions where 

hydrogen alone was unable to burn . Towards the low temperature 

ignition limits an unsteady combustion effect was cibssrved. This is 

not understood at present. but it is suggested that it might te 

connected with the selective burning of the silane fuel component at 

temperatures where the heat release is not sufficient to ignite the 

hvdrogen. 

The film cooling effect of surface injected hvdrogen was measured over 

a wider range of equivalence ratio than before. In order tcl test a 

simple model for the transfer of heat through the fuel laver to the 

wall.the injection of hvdrogen into nitrijgen and sir. helium into air 

and nitrogen into air were all tried. The simple model. which dces not 

include combustion effects, was found to be effective for qualitativ? 

comparisons between different injection conditions. Good quantitative 

correlations to specific experimental data mav be obtained tw 

adjustment of a parameter which governs the effective diffusivitv of 
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IV 

heat across the fuel laver. but it cannot at present be considered a 

true predictive technique. Values of the order 10 times that of the 

moiecuiar cnermal conductivitv are reauired. 

Heat transfer measurements without injection were repeated to ctinfirm 

previous indications of heating rates lower than simple flat plate 

predictions for laminar boundarv lavers in equilibrium flow. The 

previous results were reproduced and the discrepancies are discussed 

in terms of the model geometrv and departures of the flGw fram 

equilibrium. 

In the thrust producing mode. attempts were made to increase specific 

impulse with wall injection. This was bv means of a staged expansion. 

and. in a slight departure from two dimensional flow. transverse 

iniection from wall mounted circular Sets. Some improvements were 

evident at the lower enthalpies. The performance of wall injected 

scramiets in shock tubes still appears to tle limited bv the existence 

of a laver of cold fuel attached to the wall. 

In nrder to confirm that the pressure disturbances seen on the wall 

were also present in the duct. a pitot rake was mounted from the top 

of the model and spanned the mixing and combustig2n reginn of the fuel 

-iet. Clear evidence of transverse pressure disturbances was seen. 

Some preliminarv tests were also Derformed on shock induced ienition. 

to investigate the possibilitv in flight of iniecting fuel upstream of 

tne combustion chamber, where it could mix but not turn. Ignitian of 

the Dremixed fuel and air mixture would subsequently occur due ti:& 

shock heating at the combustion chamber intake. 

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Hydrogen was injected at a condition where it could not sustain 

combustion, and it was then heated Gv %he passage nf an 13blique shcck 

wave. In the central injection mode ~ ignition followed rapidlv after 

shock reflection. 

In the wall injection mode it was not Possible to induce ianition 

under conditions where hvdroEen would not burn without compression. 

Moreover. under conditions where combustion was possible. the passage 

of the shock did nothing to enhance combustion. The failure of shock 

stabilizsd combustion with wall injection is attributed to the 

quenching effect of the cold model walls. It is still thought that the 

technique might be successful with heated walls or with heated fuel. 
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VI 

LIST OF FIGURES 

1. Schematic of experimental apparatus 

2. Mach 6.5 configuration. 

3. Pressure profiles. ignition limit 20% SiH&/H2. M=3.4. Constant area 

duct. 0.3, T= 400 - 445K. P=24 kPa. 

4. Pressure vs time, constant area duct. 20% SiH4/H2. T=415K, P=24kPa. 

H=1.@7 MJ/kg. 

5. Pressure profiles. ignition limit 20% SiH4/H2. M=3.3. constant area 

duct, T=7@5K. P=15 kPa. H=1.94 MJ/kg. 

6. Pressure profiles. 20% SiH4/H2. M=3.3. constant area duct, T=97@K. 

P=1@ kPa. constant area duct. 

7. Pressure profiles. effect of concentration of silane at mach 4.5. 

H=3.43 MJ/kg. T=625K. P=30 kPa. constant area duct. 

9. Pressure vs time. constant pressure burning of 10% SiH4/H2 at mach 

4.5. H=2.65 MJ/kg. T=48@K, P=3@ kPa. constant area duct. 

9. Pressure profiles for constant area duct. effect of concentration 

of silane at mach 4.5, H=2.65 MJ/kg. P=3@kPa. T=48@K. 

10. Pressure prclfiles for constant area duct. effect of concentration 

of silane at mach 4.5, H=1.9@ MJ/kg. P=30kPa. T=4@@K. 

11. Pressure profiles for constant area duct. effect of concentration 

of silane at mach 5. H=8.7 MJ/kg. P=20 kPa, T=1500K. 

12. Pressure profiles for constant area duct. effect of concentration 

of silane at mach 5. H=6.2 MJ/kg. P=20 kPa, T=l@CtOK. 

13. Pressure profiles for constant area duct. effect of concentration 

of silane at mach 5. H=4.2 MJ/kg. P=2@ kPa. T=65rJK. 

14. Pressure profiles for constant area duct, effect of concentration 

of sidane at mach 5. H=3.43 MJ/kg. P=20 kPa, T=540K. 

15. Pressure profiles for constant area duct. effect of concentration 

of silane at mach 5. H=2.65 MJ/kg. P=2@ kPa. T=41@K. 

16. Pressure profile for constant area duct at mach 5. 2@% SiH4/H2 

injection. H=1.9 MJ/kg. P=2@ kPa. T=350K. 

17. Pressure profile for constant area duct at mach 5, 17.4% Ar/H2 

injection. H=4.2 MJ/kg. P=20 kPa. T=65@K. 

18. Presslire prnfiles fer ~~nstint area duct at zach 5. 17.4% Ar/H2 

injection, H=8.7 MJ/kg. P=20 kPa. T=15@@K. 

1'3. Diagram of 15 degree divergence thrust surface without cowl. 

20. Diagram of 15 degree divergence thrust surface with cowl. 
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21. Diagram of 15 degree divergence svmmetrical thrust surface 

22. Thrust contribution due to an expansion wave passing through a 

mach number gradient. 

23. Thrust contribution due to a burning fuel jet. 

24. Pitot profiles across an open duct. 15 degree divergence. M=3.5. 

P=160 kPa. T=1700 K. =1.0 - 

25. Pitot profiles across an 6pen duct. 15 degree divergence. M=3.5, 

P=160 kPa. T=ZSO@K. =1.27. 

26. Pitot profiles across a closed duct, 15 degree divergence. M=3.5. 

P=160 kPa. T=250@K, =1.0,1.85. 

27. Pitot profiles across a svmmetrical duct. 15 degree divergence. 

M=3.5. P=3.5. T=25@0K. =1.09. 

28. Central injectors. source flow and standard. 

29. Wail injectors. Source flow and standard. 

30a. Pressure profiles for central injection from a source flow and a 

cvlindrical nozzle into a constant area duct. H=4.2 MJ.kg, =1.38. 

M=3.5. 

30b. Heat transfer rates for central injection from a seurce flow and 

M=3.5. 

cvlindrical nozzle into a constant area duct. H=4.2 MJ/kg. =1.38. 

31a. Pressure profiles €or central injection from a sclurce flew with a 

M=3.5. 

cvlindrical nozzle into a constant area duct. H=8.7 MJ/kg. =l. 72. 

31b. Heat transfer rates for central injection from a source flow and 

M=3.5 

a cvlindrical nozzle into a constant area duct. H=8.7 MJ/kg. =1.72 . 

32a. Pressure profiles for wall injection from a source flow and a 

cvlindrical nozzle into a constant area duct, H=4.2 MJ/kg. =1.48. 

M=3.5'. 

32b. Heat transfer rates for wall injection from a source flow and a 

cvlindrical nozzle into a constant area duct. H=4.2 MJ/kg. =1.48. 

M=3.5 

33a. Pressure profiles for wall injection from a source flow and a 

cvlindrical nozzle into a constant area duct.H=8.7 Mi/kg. =2.0. 

M=3.5. 

33b. Heat transfer profiles fsr wall iniection from a saurce flow and 

M=3.5. .- cvlindrical nozzle into a constant area duct. H=8.7 MJ./kg. =c. 9. 

34a. Predicted and measured values of the heat transfer rate for a 

constant area duct with injector. H=8.7 MJ/kg. M=3.5. 
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VI11 

34b. Predicted and measured values of the heat transfer rate for a 15 

degree diverging duct with injector. H=8.7 MJ/kg. M=3.5 

34c. Predicted and measured values of the heat transfer rate for 2 IC; 

degree diverging duct with injector. H=4.2 MJ/kg. M=3.5 

35. Heat transfer rates for the injection of hydrogen into a nitrogen 

test gas in a constant area duct, H=4.2 MJ/kg. M=3.5. 1.6 mm throat. 

36. Heat transfer rates for the injection of hvdrogen into 3 nitrogen 

test gas in a constant area duct, H=8.7 MJ/kg. M=3.5. 0.3 mm throat. 

37. Heat transfer rates for the injection of hydrogen into a nitrclgen 

test gas in a constant area duct. H=8.7 MJ/kg. M=3.5, 1.6 mm throat. 

38. Heat transfer rates in a constant area duct when hvdrogen is 

injected into a nitrogen test gas with different injectors. H=8.7 

MJ/kg. M=3.5. 

39. Heat transfer rates in a constant area duct when hydrogen is 

injected into a nitrogen test gas with different injectors. H=4.2 

MJ/kg. M=3.5. 

40. Schematic of simple heat transfer model. 

41. Variation of- m v with m ,M=3.5 

42 . Variation of Tiv with m . M=3.5. 

43a. Pressure profiles for the injecticon of hvdrogen into a constant 

area duct. nitroeen, air test gases, H=8.7 MJ/kg. =1.35, M=3.S. 

43b. Heat transfer rates for the injection hvdrogen into a constant 

area duct, nitrogen. air test gases. H=8.7 MJ/kg. =l. 35. M=3.5. 

44a. Pressure profiles for the iniection of hvdrogen into a constant 

area duct. nitrogen, air test gases. H=8.7 MJikg, =2.60, rd=3.5. 

44b. Heat transfer rates for the injection of hvdrogen into a constant 

area duct. nitrogen. air test gases. H=5.7 MJ/kg. ~2.6, M=3. E 

45. Heat transfer rates for the injection of nitrogen intcl air and 

hydrogen into nitrogen at similar values of m v . H=8.7 MJikg. M=3.5. 

46. Proposed central injector expanded from centre line. 

47. Schematic of 2 stage expansion nozzle. 

48. Schematic of expansion wall let interacticln. 

49. ISP/ . Comparison of dual and single staze expansions, 

50. P/X. Dual stage expansion. short duct, parallel wall injection. 

M=3.5, Hs =4.2 arid 5.7 Mjikg. 
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IX 

52. P/X. Dual stage expansion. long duct. parallel wall injection. 0.9 

mm throat, Hs=8.7 Mj/kg and 4.2 Mi/kg. 

53. Schematic of transverse arid parallel injection. 

54. Isp/phi Transverse injection. Dual stage expansion. long duct. 

M=3.5. Hs=8.7.6.1 and 4.2 Mj/kg. 

55. Isp/phi. Various configurations of wall iniectian. Hs=8.7 and A.2 

Mj/kg, M=3.5. 

55. Isp/phi. Long duct. dual and single stage expansions. 27% 

transverse injection. Hs=8.7.6.l.and 4.2 Mj/kg. M=3.5. 

57. Ccimparison of transverse and parallel injection. long duct. single 

expansion, P/x and T/x. M=3.5 

58. Central injection shock stabilized cqmbustion rig. 

59. Pix. Shock stabilized combustion central injection. 5 degree 

shock. M=5. 

60. Wall injection shock stabilized combustion rig. 

61. P/X. Shock stabilized combustion wall injection. 10 degree 

c hclc k , Hs =8 .7 M j / kg . 

62. Pix. Shock stabilized combustic.n wall injection. 10 degres skier-k. 

Hvdrogen and helium injection comparisons. Hs=8.7 Mj/kg. M=5. 

63. P/X. Shock stabilized combustion wall injection. 10 degree shock. 

M=3.5. Hs=8.7 M j ikg. 

64. P/X. Shock stabilized combustion wall injection. 10 degree shock. 

M=3.5. Hs=b.l M.i/kg. 

65. P/X. Shcjck stabilized combustion wall iniection. 10 degree shock. 

M=3.5. Hs=4.2 Mj/kg. 

66. PYX. Silane injection from the wall, M=3.5. Hs=4.2 Mjikg. 0.5. 

67. Effect of freestream oxvgen radical concentration on ignition 

delay times in 20% SiH4/H2 mixtures. results of computer simulations. 

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INTRODUCTION AND CONCLUSIONS 

This constitutes the second progress report on Shock Tunnel'Studies 13f 

Scramjet Phenomena under NASA contract No. NAGW-674. 

The purpose of the study was to look more closelv at some af the 

features identified in the first progress report. Ref 1. 

The report is divided into sections as outlined below. 

Section A. Silane c.ombustion. 

Preliminarv tests in 198.5 showed the effectiveness of silane as an 

ignition aid when mixed with hvdrogen at 20% bv volume. Further 

experiments were performed using silane at different concentrations 

and over. a wider range of pressure and enthalpies. The silane was 

found to be effective at concentrations as low as 5%. 

In order to confirm that the combustion enhancement was du2 to the 

chemical kinetics of the silane. rather than a gas dvnamic effect. 

control experiments were Derformed using a mixture of argon and 

hvdrogen. The concentration was adjusted to produce a mixture with the 

same molecular weight as the 20% silane mix. It was found that this 

mixture was unable to burn at conditions where the silane mixture 

burned vigorouslv. This was taken as confirmation of the chemical 

nature of the silane ignition mechanism. 

In studies of wall injection the existence of a quenched laver 

attached to the wall was postulated to explain the low values of 

specific impulse, and also the lack of combustion at low equivalence 

ratios. Silane provides a convenient means of checking this theorv. as 
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it is capable of spontaneous combustion at the wall temperature. A 20% 

silane mixture was injected from the wall into a constant area duct at 

an equivalence ratio which wouia nor support combustion with hvdrogen 

alone. Large pressure rises were recorded down the duct. which is 

consistent with the existence of a laver of mixed air and fuel which 

is attached to the wall. but unable to burn due to its' lclw 

temperature. This suggests that shock tunnel experiments with cold 

wall might be better able to simulate flight conditions with hot walls 

bv the use of small concentrations of silane. At present attempts to 

investigate modifications. such as transverse wall injection and 

staged expansions. are restricted because the presence of the wall 

quenched zone is the Predominant factor in determining the amount of 

heat release and thrust. 

Section B. Pitot rake measurements. 

Static pressure measurements have shown the existent? of disturbances 

along the wall and the thrust surface which are consistent with the 

simple thcorv developed for the operation of a two dimensional 

scramlet. These disturbances would be expected to reflect off the 

walls and propagate across the duct. creating significant transverse 

pressure gradients. A pitot rake was constructed to measure the 

transverse pitot pressure profiles behind a normal bow shock. and this 

was Gsed in several thrust producing configurations. As expected. 

transverse pressure disturbances were observed. 
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D 

Section A. Source flow nozzles. 

The fuel injection nozzles used in the experiments are of simple 

ccjnstruction with sudden changes of geometrv. and in some cases a lack 

of symmetry. Equivalence ratio control is bv means of adiusting 

hvdrogen reservoir pressure, and differences in mach number and jet 

thickness occur between shots. The nozzles are not correctlv expanded 

for all conditions. and freestream pr5ssure is established thi-~~uch a 

series of shocks and expansions. This is necessary in order to 

maintaxn a reasonable run rate in the limited tunnel time available. 

It is difficult to model the flow through the 

the effects of the variations between conditions are not known. 

Control experiments were therefore performed with source flow nozzles 

for both wall and centr31 injection. The nozzles were designed with 

the same throat area as the standard nozzles. but with uniform 

diverging sections in the supersonic regions. While these nozzles were 

still not operated at perfectlv gxpanded conditions. thev did provide 

comparisons for the effects of internal nozzle geometrv on otherwise 

identical runs. 

nozzles accurately, and 

No difference could be observed between the nozzles from either 

pressure or heat transfer measurements. 

Section D. Fuel off tmundarv layer heat transfer measurements. 

Fuel off heat transfer measu-rements with a wall injector fitted taken 

during the test program in 1985 showed surprisinglv law values of heat 

transfer when compared to flat plate predictions. The difference 
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between this and previous heat transfer work was that the presence of 

the wall injector provided a step which caused the boundarv laver to 

temporariiv separate. 

This work was repeated with new instrumentation tu confirm the result, 

and comparisons were also made with experiments bv previous workers 01-1 

different models. The same results were again obtained, and the 

discrepancies are attributed to non equilibrium flow effects. and also 

to' the fact that the model was not a flat plate and flat plate 

predictions would not be expected to apply preciselv. 

Section E. Film cooling studies. 

The effectiveness of wall injected hvdrogen as a coolant was 

demonstrated in 1985. The data obtained was mainly at high equivalence 

ratios. and no theoretical treatment was offered to explain the 

results. More work is reported here giving mere data at low 

equivalence ratios. and a simple predictive technique is presented. 

The model gives a good qualitative explanation of the results. and 

identifies that the critical oarameter for assessing the cooling 

effect of a non reacting insulating layer is the prclduct of mass flow 

rate with the thickness of the let. 
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Section F. Two stage divergence and transverse iniection. 

The first experiments with wall iniectican were performed in 1965. and 

for theoretical reasons it was expected that an improvement in 

performance could be produced bv means of a multistaged expansion and 

transverse circular jets of fuel injected upstream of the parallel 

wall component. 

Both techniques were to a certain extent successful. but 

interpretation of the results is complicated bv the presence of the 

wall quenching lav2r. 

The two stage expansion was successful in encouraging combustion in an 

expanding section which is sirnultaneouslv developing thrust. but tie 

thrust obtained on further expansian of the flow was disappointing. 

Transverse injection of the fuel was seen to give ignition at slightlv 

lower equivalence ratios than parallel injection. but this improvement 

in performance was not maintained when the equivalence ratio was 

incrqased. 

Section G. Shock Induced Ignition 

Weight minimization of scramjets is just as important as thermodvnamic 

Performance if a flight engine is to be produced. dne means of 

rsducing weight is to inject the fuel upstream of the intal.e, where it 

can mix but it is not hclt enough to turn. Ignition can later be 

initiated bv compressions at the combustion chamber intake. 
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This effect was examined for both wall and central injection bv means 

of an oblique shock wave passing through the duct. Fclr the central 

iniector rapid combustion of the premixed fuel occurred after the 

shock at freestream conditions that would not normally support 

cornbustion. 

With wall iniection no such effect was observed, and at conditions 

where combustion was possible no improvement was Droduced bv the shock 

heating. This is again attributed to the presence of a quenched layer 

attached to the wall. and emphasises the need for heated walls or 

heated fuel for comprehensive studies of any combustion effects with 

wall iniection. 

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SECTION A 

COMBUSTION LIMITS OF SILANE 

Silane was mixed with hydrogen at different molar ratios and was 

centrallv injected into a constant area duct. Silane concentraticm was 

varied bv adding additional hydrogen to a pre-prepared 20% 

siiane/hvdrogen mixture . Convection currents caused bv a heating coil 

secured to the bottom of the mixing cylinder were used to mix the 

gases. Work was carried out at three Inach numbers and intake 

pressures. Temperatures were reduced for each case bv reducing the 

enthalpv of the air flow, thus vielding a range of ignition 

conditions. Mach 3.4 and 5 contoured nozzles were used. Mach 4.5 

flows were produced bv using a model intake attachment in conjunctian 

with the mach 5 nozzle. This attachment was in the form of two plates 

inclined at 2.57 degrees as shown in figure 2. 

COMBUSTION LIMIT OF 20% SILANE/H2 MIXTURES AT MACH 3.4 

Work ,at this mach number for determining the combustion limits of 

silane /hydrogen mixtures was performed at the Universitv of 

Queensland using shock tunnel TQ . 

A 20% mole fraction mixture of silane in hvdregeis was centrallv 

injected into a constant area duct with flews of a nominal mach number 

af 3.4. Figure 2 .with the compression wedges removed. represents thT 
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above configuration. Figure 3 shows that with a static intake 

pressure of 24 kPa that the temperature must be reduced to around 400K 

before the silane mixture is extinguished. 

It should be noted that the temperatures listed in the figure are 

nominal, but thev do indicate the relative magnitudes of. the 

temperature changes involved. 

At 4lSK the combustion becomes unstable and the pressure rise shown in 

figure 3 is due in fact to a pressure rise propagating upstream with 

time. -It is suggested that this might be due to a thermal choking 

effect associated with an increased ignition delay time which would 

prevent combustion until the air/fuel mixture has travelled the length 

of the duct. Once combustion starts at the end of the duct. the 

heat release is such that steady flow cannot be maintained, and a 

shock is created which travels upstream. Figure 4 illustrates this 

effect with pressure against time traces at stations along the length 

of the duct. An unsteadv effect was observed near the combustion 

limits for silane/H2 mixtures at all mach numbers and concentrations 

of silane . 

Experiments at mach 3.3 show that 20% silane mixtures are extinguished 

at 705K when the pressure is reduced to 15 kPa (see figure 5I.However. 

vigorous burning is observed at an even lower pressure of 10 kPa and 

a temperature of 970K (figure 6). Experiments at intermediate 

cclnditions to establish a smaller range of temperatures where silane 

is shown not to burn are planned. . but the present results give a clear 

indication of the sensitivity of the combustion temperature limit with 

pressure at these low pressures. 

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COMBUSTION LIMITS OF 20%. lo%, AND 5% SILANE / HYDROGEN MIXTURES 

AT FLOWS OF MACH 4.5 

Work at this mach number was performed at the A.N.U. in Canberra using 

shock tunnel T3. 

Again the mixture was iniected into a constant area duct configuration 

using central in3ection. Experiments wwe performed at three 

enthalpies namely 3.43. 2.65. and 1.3 MJ/kg which correspond to an 

intake pressure of 30 kPa and temperatures of 625K. 480K. and 400K 

respec tivelv. 

At 30 kPa and 625K (3.43 MJ/kg), the 20 ”/. and 10% mixtures burned 

with about the same amount of heat release. but the 5% mixture was 

apparently extinguished (at least over the length of the duct which 

was considered). This is shown in figure 7 in plots of duct static 

pressure normalized with stagnation pressure plotted against distance 

along the duct. The 20% mixture appears to burn with a marginallv 

shorter ignition delay time than the 10% mixture. 

At 30kPa and 480K (2.65 MJ/kg) all of the mixtures shew an unusual 

pattern of tlehaviour. The pressures along the length of the duct rise 

up to a particular level. and then stav roughlv constant descjite the 

falling stagnation pressures. This is illustratsd in figure S where 

the pressures at stations 3.4.5 and 6 are Plotted against test timi: 

for the 10% mixture. When nclrmalized against stagnation pressure. 

duct static pressure vs distance traces appear unsteadv. so for this 
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condition plots of normalized pressures are taken at comparable 

stagnation pressure (figure 9,. These show the 10% mixture burning 

with the greatest pressure rise followed bv the 5% mixture then the 

20% mixture. This effect is not understood but appears to clCCur onlv 

at these lower temperatures. The fact that there appears to be 

unsteadv burning of the 5% mixture at this temperature desoite no 

combustion at the higher temperature cannot be explained at this 

stage. 

At 30 kPa and 400K (1.9 MJ/kg). the same unsteadv effect was cltlserved 

with plots of normalized static pressures vs distance along the duct 

(figure 10) showing the 5% mixture giving the greatest pressure rise 

followed bv the 10% then 20% mixtures. The decrease in heat release 

with silane concentration suggests that the pressure rises are not (due 

to the selective burning of the silane component of the mixture and 

that the hvdrogen is somehow being burnt as well at these low 

temperatures. 

The mach 4.5 condition was obtained bv use of a scoop intake (figure 

2) and in this series of tests. it proved difficult to achieve 

perfectlv steady normalized pressure traces. This may be the cause of 

the apparentlv anomalous result produced at mach 4.5. 
11 

COMBUSTION LIMITS OF 20X, 102. 5X AND 2-57! SILANE/HYDROGEN MIXTURES AT 

MACH 5 

A silane /hydrogen mixture was iniected into a mach 5 air flrlw at 

equivalence ratios from 0.7 to 1.3. The effects of varying the 

silane concentration was investigated over a range of enthalpies.. 

In.the enthalpv range 8.7 to 3.43 M~/kg. which corresponds to intake 

temperatures in the range 1500 to 540 K. steady cclmbusticin was 

observed. Hvdrogen fuel alone had previouslv been found not to turn at 

these conditions. At an enthalpv of 2.65 Mj/kg a combustion effect was 

observed. but it was not steadv. 

This is an effect which has been observed with silane over a range of 

conditions as the temperature approaches the lower limits for 

combustinn. A possible explanation for this effect would be that the 

silane is burning. but the temperature rise is insufficient to ignite 

the hvdrogen. The scramjet would therefore see a low equivalence 

ratio silane fuel jet which would probablv be completely burned and 

produce constant heat release as the intake pressure drops. This 

effect would tend to hold the internal duct pressures high as the 

intake pressure dropped. leading to unsteadv duct Pressures when 

normalized to the intake or stagnation pressures. In this unsteadv 

region the duct pressure rises would be expected to increase with 

silane concentration, and this was observed in some cases. 

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This is in contradiction with the data obtained at mach 4.5. and 

suggests that there may be two mechanisms for unste2dv combustion at 

low enthalpies. The first. consisting Gf a thermal choking effect. and 

the second being selective combustion of the silane alone. 

At 8.7 Mi/kg silane was injected at concentrati@ns of 5% and 20%. and 

no significant difference between the two was observed. This is shown 

in Figure 11 and suggests that at the higher temperatures that 

raising the level of silane concentration has onlv marginal benefit. 

At 6.1- Mj/kg silane was injected at 2.5% and 20% and the same steadv 

pressures were achieved down the duct, but with a slightly longer 

delav in the 2.5% case. This is illustrated in figure 12. 

At 4.2 Mj/kg silane was injected at 2.5. 5. 10 and 20% concentrations. 

and the effect on delav times was seen to change stronglv between 2.5 

and 5'7, but not much between 5% and 212% (Figure 13). The steadv 

pr2ssure levels downstream in the duct were again seen to be fairlv 

independent of silane concentration. 

At 3.A3 Mj/kg all concentrations showed a significantly increased 

ignition delay (Figure 14). with not much variation bzttween the 

concentrations. However in the dewn stream regions of the duct. higher 

pressures were achieved with the lower silane concentrations. 

At 2.65 Mj/kg the combustion effects were not steadv, as mentioned 

above and the higher concentrations of silane were seen to produce, 

higher pressures. This is shown in Figure 15. 
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At 1.9 Mj/kg silane injected at a concentration of 20% did not produce 

any ignition (Figure 16) and no further work was done at this 

condition. 

The mach 5 tests produced very steady time resolved traces. and thr 

results are considered to be more reliable than those presmted for 

mach 4.5. 

In conclusion to this section. it is noted at the higher enthalpies 5% 

silane is adequate for ignition. and no improvement is produced for 

higher.concentrations. At lower enthalpies, combustion proceeds faster 

but with less heat release, as the silane concentration is increased. 

The unsteady effects at the lower temperature ignition limit, while an 

interesting phenomena. cannot be explained completelv at this stage. 

This is not considered to be critical because in practice. flight 

conditions which promote unsteadv burning will be avoided. However. it 

will be important to define the limits where these effects are 

observed. 

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14 

HYDROGEN / ARGON MIXTURES 

A Hydrogen/Argon mixture equivalent in molecular weight to a 20% 

silane in hvdrogen mixture was centrallv injected into a const-ant 

area duct. Fuel was injected such that the mass flow rate of the 

argon/hvdrogen mixture was equal to the mass flow rate of a 30% 

silane/hvdrogen mixture injected at an equivalence ratio of about one 

based on the hvdrogen content of the fuel. Flow conditions which did 

not support hvdrogen combustion but did support combustion of a 20% 

silaneihvdrogen mixture were used. This was done to test whether the 

increased molecular weight over pure hvdrogen fuel was a contributing 

factor-in the mechanism of burning when silane was added to hvdroyen. 

Figure 17 shows pressure profiles for fuel on and fuel off 

cases with a hvdrogen/argon mixture at a flow conditisn of mach 5 and 

a freestream ent-halpv of 4.2 MJ/kg. It is clearly evident that there 

is no burning. 

The freestream temperatur2 was increased bv increasing the flow 

enthalpv to 8.7 MJ/kg and still no combustion was abserved. This is 

shown in figure 18. It was concluded therefore that increased 

molecular weight of a silane/ hvdrogen mixture over pure hvdrogen 

had na significant effect on the reactivity of the fuel. Ftather. 

burning is due to the chemical kinetic mechanisms related %o the 

breakdown of silane. 

Recent numerical analysis has shown that the ignition of silane is 

verv sensitive to the presence of oxvgen zitoms created in the shctck 

tube and frozen in th? freestream flow. This effect is discussed in 

the appendix and represents an imwrtant difference between shock 

tunnel simulations and real flight conditions. 
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WALL INJECTION OF SILANE. 

The presence of the cold model walls represents a sink of heat to ths 

flow, because in the short duration of the tests the wall temperature 

does not rise significantlv above ambient. The cooling effect of the 

wall penetrates a significant distance into the flow. With cmtral 

injection this does not have a critical effect on the development of 

combustion. 

However. when the fuel is injected from the wall. there is alwavs a 

region-whose temperature will be held below the ignition temperature. 

regardless of how much combustion may take place further away from the 

wall. The hvdrogen contained in this low temperature region appears to 

correspond to an equivalence ratio of approximately 1.5. becauss no 

ignition at all is possible at lctwer equivalence ratios. This 

represents a serious defect in the ability of shock tunnel tests to 

accuratelv model a real flight situation with aerodvnamitallv heated 

walls. 

The fuel in the quenched region may be mixed with oxygen. with onlv 

the lpw temperature inhibiting combustion. To confirm that this is 

indeed the case, and that the lack of combustion is not the result of 

some other cause, a test was done with the injection of a 20% silane 

hydrogen mixture. 
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The fuel was injected at an equivalence ratio of 0.5 into a flow of 

enthalpy 4.2 Mj/kg. With hydrogen injection alone no combustion would 

be expected at this condition. The results of this test are shown in 

Fig 66, and it can be seen bv the pressure rise above the fuel off 

levels that it burnt well. 

This result is significant in that it demonstrates that oxygen is 

diffusing to the fuel laver close to the wall. and that it is thermal 

effects which are preventing it from burning. This gives encouragement 

to efforts which are currentlv being made to design a model which can 

use heated fuel, and possibly even heated walls. It also suggests that 

in the absence of the above, silane might be used to investigate other 

aspects of combustion which are difficult to study properly in the 

presence of an extensive laver of quenched fuel. This includes the use 

of staged expansions and transverse iniection discussed in section F. 
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SECTION B 

PRESSURE PROFILES ACROSS A DIVERGING DUCT 

For these experiments. a "pitot rake" was secured to the model to 

take a pressure profile normal to the flow direction. Three cases at 

mach 3.5 with central fuel injection were tried, 15 degree divergence 

thrust surface without cowl (fig. 19). 15 degree divergence thrust 

surface with cowl (fig. 20). and twin 15 degree divergence thrust 

surfaces (fig. 21). Measurements were those of stagnation pressure 

behind-bow shocks created bv the presence of the blunt pitot probes. 

Pressure disturbances measured on the walls of the thrust surface 

indicate the presence of compression and sxpansion waves in the flow 

that are consistent with a simple model that has been developed for 

the combustion and thrust prclduction mechanisms in a two dimensinnal 

scramiet iref. 2). These waves will reflect off the walls and should 

produce corresponding transverse pressure gradients across the duct. 

This series of experiments has been developed to confirm the presence 

of such transverse disturbances and add further credibilitv to the 

theoretical model. A full treatment to predict the quantitative 

transverse pressure profiles for these test cases has not vet teen 

performed. but the results are still significant in confirming the 

presence of such effects. 
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EXPLANATION OF THE MODEL 

Heat release from the fuel jet results in a subsequent drop in mach 

number in the jet because of increased temperature. Because of the 

existence of a negative gradient of mach number from the edge of the 

burning jet to the middle, the expansion waves produced bv the 

divergence, create reflected compression waves and then expansion 

waves as thev progress through the burning jet. This is illustrated 

in figure 22, which shows the passage of a single expansion wave 

through a mach number gradient. Integration of the effects of all 

the expansion waves in the expansion fan passing through the fuel iet 

produces a net pressure rise on the thrust surface. Compression and 

expansion waves reflect off solid surfaces and continue on to the rear 

of the duct where thev are sensed bv the pitclt rake. 

Coupled with this effect is the Production of compression waves due 

to the mixing and burning of the injected fuel. and their subsequent 

reflection from the thrust surface as shown in Figure 23. The 

compression are later followed bv expansion waves from the combustion 

region, but the net result in a confined duct is a pressure and 

thrust increase. 

Wave propagation paths will be different for each of the different 

configurations and each is discussed separatelv. together with 

experimental results. 

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To carrv out a full analysis . mach number profiles for the burning 

hydrogen jet would be obtained from a two dimensional mixing and 

combustion program (8). Integral expressions for the strength of 

reflected expansion and compression waves from the mach number 

gradient could then be evaluated to give pressure profiles (ref 2). 

OPEN DUCT - 15 DEGREES DIVERGENCE 

When the scramiet is configured as a closed duct. expansion waves 

propagating from the end of the shock tunnel nozzle clear the intake 

of the-scramjet and have no effect on the thrust surface pressure 

measurements. However. for this configuration with onlv one'surface.to 

ensure that expansion waves from the nozzle exit did not impinge on 

the instrumented part of the thrust surface, the nozzle was extended 

bv a sector which fitted between %he side plates of the scramjet. 

As seen in figure 19. the first expansion wave encountered bv the 

thrust surface from the edge of the nozzle extension is well 

downstream of the surface pressure transducers. Note that expansions 

propagating from the edge of the nozzle are bent 15 degrees in 

towards the thrust surface by the presence of the expansion fan at the 

corner of the diverging section. The pitot rake experiences these 

expansions due to the lack of a top cowl. 

Plots of pitot pressure versus v (distance from datum Ai are srirJwn in 

figures 24 and 25 for fuel off and fuel on cases at two freestream 

enthalpies. 6.1 and 8.7 MJ/Kg:- It appears from these plots that 

changing the enthalpy of the air flow has marginal effect on the 

stagnation pressures across the duct. 
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CLOSED DUCT - 15 DEGREES DIVERGENCE 

With a closed duct. expansions created bv the start of divergerice 

reflect off the top cowl and continue on down the duct. Coupled with 

this effect is that of the compressions and expansions reflected from 

the mach number gradient caused bv the presence of the burning fuel 

jet. Overall pitot pressures are lower than those for the case when 

the duct has no cowl. This is probablv due to multiple reflections 

of the expansion caused bv the start of divergence. For this case. the 

pitot rake was traversed up and down the width of the duct for the 

fuel on case with equivalence ratio of one. Hvdrogen at a higher 

equivalence ratio was injected and had little effect on the 

pressure profile. Plots of fuel on and fuel off pressure profiles at a 

freestream enthalpv of 8.7 MJ/kg are shown in figure 26. 

SYMMETRICAL DUCT - 15 DEGREES DIVERGENCE 

When the scramiet was configured as a svmmetrical duct. there was a 

mGre marked difference between fuel on and fuel off results CGmpared 

with the closed duct case. Plots of fuel off and fuel on pressure 

profi.les at a freestream enthalpv of 8.7 MJ/Kg are shown in figure 27. 

COMMENT 

Although the results of the pi%ot survey cannot De fully interpreted 

until the computaticns mentifined abave had been performed. thev d@ 

indicate the presence of substantial transverse pressure gradients. 

This is consistent with the thrust production mechanisms pos%ulated 

in reference 121. 
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21. 

SECTION C 

SOURCE FLOW NOZZLE 

Fuel nozzles used in the experiments to date have been designed to give a wide 

range of equivalence ratios with the simplest possible design and a minimum of 

structural adjustments between operating conditions. This was done to make 

optimum use of the limited tunnel test time which is available to the project. 

The nozzle consists of a removable two dimensional throat, and the mass flow 

rate and Mach Number are controlled by the setting of the fuel reservoir 

pressure. Figures 28a and 29a show the standard nozzle configurations for 

wall and central injection. Ref (1) Figure 3 shows the range of injection 

conditions which may be achieved with the wall injectors. For a given 

injection condition the nozzle exit pressure will not in general be matched to 

the free stream pressure, and a series of compressions and expansions will be 

created as the pressures equalize. It is not known what effect an under or 

over expanded jet will have on the mixing and combustion rates of the 

scramjet, and this process is difficult to model numerically with the abrupt 

changes in geometry of the standard injector. 

However previous experience with wall injection suggests that it is the 

equivalence ratio alone which controls the amount of mixing and combustion. 

Ref (1) presents data in Figures 8 and 9 comparing runs with similar 

equivalence ratios, but obtained with different throats which produce different 

exit pressures. No significant difference is apparent between the two 

injectors for either of the enthalpies considered. 

In order to further investigate this effect nozzles with the same throat sizes, 
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22. 

but with different geometries were constructed. The geometry of the new 

nozzles was chosen so that it would be more ameanable te n1zr.erical azalysis, 8s 

well as giving a direct comparison with the standard injectors. 

A comparison was made between two different injectors for both central and wall 

injection. The two injectors used were a cylindrical nozzle and a source-flow 

nozzle. Figures 28b and 29b show the different structures of the two injectors 

for central and well injection. Both injectors are two dimensional and expand 

the fuel from a 0.9 mm throat (wall injection) and a 1.6 mm throat (central 

injection) to 5 mm at the nozzle exit. The source-flow nozzle expands the fuel 

so that the flow through the nozzle is similar to that from a two-dimensional 

line mass source. 

For both central and wall injection pressure and heat transfer measurements 

were taken simultaneously. The freestream test gas was air and the stagnation 

enthalpies of the test gas were either 4.2 mJ/kg or 8.7 mJ/kg. Hydrogen fuel 

was injected from the wall injector at equivalence ratios of 1.45 and 2.0 into 

a test gas of stagnation enthalpy 4.2 mJ/kg respectively and from the central 

injector at equivalence ratios of 1.4 and 1.7 respectively. 

The results of the central injection experiments are given in Figures 30a, 30b, 

31a and 31b and the wall injection results are given in Figures 32a, 32b, 33a 

and 33b. From the normalized pressure results it is seen that for both the 

wall and central injectors and at either test condition, little difference is 

produced by the different injectors. The heat transfer measurements also 

reflect this result, however they are not as conclusive due to the considerable 

scatter observed. 

In conclusion, no difference between the results for a source-flow nozzle and 
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23. I 

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a cylindrical nozzle could be detected within the experimental accuracy. 

Coupled with previous data for the effect of throat size at a given 

equivalence ratio, it still seems that the equivalence ratio is the factor of 

most significance in determining the mixing and combustion rates. 

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24. , 

SECTION D 

BOUNDARY LAYER HEAT TRANSFER 

Previously reported work, Ref (11, measured heat transfer rates to the scramjet 

walls at lower levels than would be predicted by flat plate predictions. A 

series of tests was performed with different heat transfer gauges to confirm 

the reproducability of these results. 

0 Heat transfer measurements to the wall of a constant area duct and a 15 

diverging duct, both with 5 mm steps, were taken simultaneously with pressure 

measurements. Experiments were done at an intake Mach number of 3.5 and 

enthalpies of 4.2 and 8.7 MJ/kg. From the pressure measurements and the intake 

conditions emperical predictions of the heat transfer rate are made using the 

correlations of Ref 3 for a flat plate laminar boundary layer and following the 

treatment of Ref 4 for a turbulent boundary layer. The intake conditions are 

determined from the stagnation conditions using a computer model for the 

non-equilibrium expansion of a reacting gas down the nozzle, Ref 5. 

Similar measurements have been taken previously and are recorded in Ref 1. 

For those experiments the measured heat transfer rate was of order 1/10 of the 

predicted value for a turbulent boundary layer and was less than or equal to 

the predicted value for a laminar flow. Moreover, there was a greater 

deviation between the measured and predicted laminar values when more 

expansions were experienced by the flow, and this deviation was more 

pronounced at higher enthalpies. 

The results of the latest experiments and predictions for a laminar boundary 

layer are given in Figures 34a, 34b and 34c. The turbulent boundary layer 
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predictions were again of order ten times those measured and from these figures 

it is concluded that the boundary layer is laminar. It can be seen that the 

same trends were observed again. It is also seen that the predicted values 

approach the measured values at the downstream end of the duct. A possible 

explanation for the low measured values lies in the non equilibrium chemistry 

of the flow at the higher enthalpies. 

The test gas in a reflected shock tunnel is produced by expanding the gas from 

a high pressure and temperature reservoir which is in equilibrium, and, due to 

the high temperatures, significant amounts of dissociation occur, particularly 

of oxygen molecules. During the expansion the dissociated products recombine 

at a finite rate, and if the nozzle is sufficiently short equilibrium 

composition will not be maintained. This leads to a reduction of temperature, 

and hence heat transfer rate, as the chemical energy which would be released 

on recombination is stored. This effect increases with stagnation enthalpy, 

and can be shown to be significant at the conditions relevant to this study, 

Ref. 6. 

As the result of expansion from the trailing edge of the injector and from the 

start of the divergent section further departures from equilibrium will occur, 

and further lowering of the heating rates might be expected. Further 

downstream the condition is nearer to equilibrium and therefore the measured 

and predicted values approach each other. This effect has been observed 

before in the same facility, Ref 6, and is more significant at higher 

stagnation enthalpies. 

Another possible cause of the lower than expected heating rates in the 

diverging duct might be in the interaction of the corner expansion with the 

boundary layer on the side walls. The effect of the expansion in the low mach 
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26. 

number boundary layer would be to create a cross flow along the wall, which 

might spill coded boundary layer gases on to the thrust surface and redi-lre the 

heat transfer rate. 
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27. 

SECTION E 

FILM COOLING STUDIES 

An initial series of experiments was reported in Ref 1 to investigate the use 

of wall injected hydrogen as a combustion chamber coolant. The study 

demonstrated the effectiveness of the fuel layer in shielding the model walls 

from the heat of the flow. New results are presented in this report to give a 

more comprehensive set of data for the testing of theoretical models. 

Results are given for the following three conditions: 

(1) Injection of hydrogen into a nitrogen test gas, 

(2) Injection of hydrogen into an air test gas, 

(3) Injection of nitrogen into an air test gas. 

All experiments used a constant area duct with the hydrogen or nitrogen being 

injected from behind a 5 mm step. Different equivalence ratios were obtained 

by varying the reservoir pressure of hydrogen and using different injectors. 

1. Hydrogen injected into a nitrogen test gas. 

Those experiments were done to model the cooling effect of injecting hydrogen 

into a test gas in which combustion does not occur. Nitrogen was chosen as the 

test gas as its molecular weight is close to that of air. Partial mixing Of 

the two gases occurs and heating of the wall results from heat being 

transported from the test gas by the mixing process in addition to thermal 

diffusion through the mixed and unmixed layers of hydrogen. Axial convection 

also transports heat downstream. 
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In this section, due to the analogy which is drawn between nitrogen and air, 

the term "equivalence ratiog1 is lis& 2s ar? slternztive Eeasuie of the i~ias~ flow 

rate of hydrogen. of hydrogen at an equivalence ratio of one is 

to be interpreted as injecting hydrogen with the mass flow rate which is 

required for a stoichiometric mixture with air. 

The injection 

The heat transfer results for the test gas at enthalpies of 4.2 MJ/kg and 

8.7 MJ/kg are given in Figures 35, 36 and 37. It is seen that for a given 

injector the increase in equivalence ratio produces an increased cooling 

effect. This results because the increase in mass flow rate (equivalence 

ratio) increases both the thickness of the hydrogen layer and the velocity of 

the hydrogen in this layer. Thus heat takes longer to diffuse through the 

thicker hydrogen layer and in addition is convected more rapidly downstream. 

These effects reinforce each other so that the heat reaches the wall further 

downstream. 

When the fuel is not injected and the stagnation enthalpy was 8.7 MJ/kg the 

heat transfer rate at the downstream end of the duct was approximately 

80 W/cm. It is also seen from Figure 36 that far downstream the heat transfer 

rate when fuel is injected approaches the fuel off heat levels. 

In Figure 38 the heat transfer rates are displayed for the injection of 

hydrogen from a 0.3 mm and 0.9 mm nozzle at equivalence ratios of 1.39 and 

1.13 respectively. It was seen from Figures 35, 36 and 37 that the heat 

transfer rate is dependent on equivalence ratio. However, from Figure 38 it 

can be seen that the Mach number is-also affecting the heat transfer rate as 

the same value of m y for different injectors is equivalent to 

00 

a different mach number. 
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29. 

A simple model for the heating of the thrust surface is to assume that a 

portion of the hydrogen is trzpped in a layer of coiistaiit thickness aiong this 

surface; see Figure 40. If it is assumed that heat diffuses across this 

boundary layer and is converted downstream and if gas dynamic effects are 

ignored so that the heat equation 

can be used, then using Laplace transforms it can be shown that the flux of 

heat into the thrust surface is 

where the length is scaled with respect to 

h+2 

cp k+l 

*lil =puy (3) moyo* 0 000 

L=--a k 

The velocity profile of the boundary layer is given by pu = pouo [+Ihwhere Y 0 

is the thickness of the injected layer when expanded to the local free stream 

pressure. It is assumed that the fuel expands isentropically from its 

reservoir pressure to the intake pressure with a mass flow rate which is 

determined experimentally as a function of reservoir pressure for each 

nozzle. The quantity AT is the difference between room temperature and the 

temperature of the isentropically expanded fuel. This term is present because 

Of the boundary condition at x = 0. This boundary condition is inconsistent 

with the velocity profile (except if k = 2) and only serves to approximate the 

boundary condition so that an analytical solution can be obtained. 
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30. 

Equation 1 assumes that transverse heat transfer is solely by means of thermal 

conduct ion. It is in fact augmented by turbulent transport and by the 

diffusion of species which may or may not be reacting. However, if it is 

assumed that an equation of the form of equation 1 may still be used, but with 

a revised value of conductivity to account for the extra heat transfer 

mechanisms, then a solution of the form of equation 2 will still apply. In 

Ref 7 the value of the effective conductivity was adjusted and good agreement 

was obtained with the experimental results. However a value of the order ten 

times that of the standard thermal conductivity was needed. As we have at 

present no independent scientific method for predicting the rate of heat 

transfer across the mixing layer, the analysis will be used in this report 

only to provide quantitative comparisons between different conditions. 

There are two distinct variables kayo and - AT , which can be varied using 

different reservoir pressures and nozzles. In Figures 41 and 42 h y and 

AT for the three injectors used and 

YO 

for a test gas at enthalpies of 4.2 MJ/kg and 8.7 MJ/kg. From Figure 42 it 

is seen that for a given injector is effectively constant for sufficiently 

large values of m . Hence as an increase in m for a given injector increases 

moyo (Figure 41) then from equation 2 an increase in m results in increased 

cooling the length scale is approximately a linear function of the mass flow 

YO 

00 

- are respectively drawn a functions of m 

0 

- AT 

YO 

0 0 

0 

(i-e. the thickness of the layer is approximately constant). 

In Figure 38 the heat transfer rates are given €or values of m y Of 

4.28 x kg/sec (Q = 1.39) through at 0.3mm throat and 4.75 x 10 kg/sec 

(Q = 8 x 1.13) through a 0.9 nun throat. It is seen that the heating rates are 

similar. An explanation for this can be derived from equation (2). The Value 

00 

-4 
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31. 

of qo is assumed to be the fuel off heat flux to the wall, which for a test 

The ..-.l..- -5 LL^ 

conductivity constant k for hydrogen at 300 K and 1 atm is 0.18 W/mK, but due 

2 

gas of enthalpy of 8.7 Mj/kq is approximately 80 w/c . "rn&UC VL LLLC 

m 

to the turbulent transport and the variations of k with temperature, a precise 

value of k is unknown. However, even if k were to increase by a factor of 

five and a was not too small it is seen from Figure 42 that would still only 

be a second order correction to q . Thus if two injectors of different throat 

0 

size were to inject hydrogen so that m y is the same for both then the heat 

flux into the wall should be similar. 

00 

A result contrary to Figure 38 is given in Figure 39 which displays the heat 

transfer rates for injection through the 0.9 mm and 1.6 mm throat with values 

of m y of 1.3 x 10 kg/sec (@ = 1.93), and 2.5 10 kg/sec (@ = 2.49) 

respectively. The enthalpy of the test gas was 4.2 MJ/kg. Therefore under 

-3 -3 

00 . 

the above assumptions the heat transfer rates should not be the same. The 

- -1 

will be less accurate 40 assumption that - AT Ka is of second order to 

for the lower enthalpy test gas. Furthermore if fi, or equivalence ratio, is 

the same for the two injectors then it is seen from Figure 42 that by using 

larger throat the value of - AT decreases and thus by equation 2 the heat 

transfer rate is increased. Hence if this effect were no longer second order, 

yo 

then for the above mass fluxes the cooling effect created by the larger value 

of m y using the 1.6 mm throat may be offset by the heating effect created 

by using the larger throat. 

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32. 

2. Hydrogen injected into-an air test gas. 

In Figures 43a, 43b, 44a and 44b the pressure and heat transfer rates recorded 

for the injection of hydrogen into an air test gas are compared with injection 

into a nitrogen test gas. Both gases had an enthalpy of 8.7 MJ/kg. The 

hydrogen was injected through the 0.3 mm and 0.9 mm throat at equivalence 

ratios of 1.34 and 2.7 respectively. It was reported in Ref l.that burning 

started at an equivalence ratio of 1.3. At lower equivalence ratios it is 

believed that the close proximity of the wall quenches the burning. This is 

confirmed in Figure 43a where it is seen that only a slight increase in 

pressure is observed for an equivalence ratio of 1.35. Furthermore this 

pressure rise is significant only towards the end of the duct. Additional 

confirmation is obtained from the heat transfer results in Figure 43b where 

it is seen that the heat transfer rate from the nitrogen test gas is the same 

as from the air. 

From Figure 44a it is seen that pressure rises occur everywhere downstream of 

the injector when hydrogen is injected at an equivalence ratio of 2.6. Thus 

some of the fuel has been burnt. It is seen in Figure 44b that in the 

downstream position of the duct the heat transfer rate for an air test gas 

appears to be higher than a nitrogen test gas, which is assumed to be the 

result of burning. It is observed that a significant pressure rise does not 

occur until halfway down the duct, and yet the heat transfer rate increases 

only in the last quarter of the duct. This delay in the heating rate also 

supports the theory that burning will not occur at the wall, for if it did 

there would be no delay time. The delay time is caused by the convection 

process transporting heat downstream while it diffuses across the- unburnt 

layer of hydrogen. 
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33. 

3. Nitrogen injected into an air test gas. 

The heat transfer rates when nitrogen is injected into an air test gas at 8.7 

MJ/kg and when hydrogen is injected into a nitrogen test gas at the same 

enthalpy are given in Figure 45. These measurements have been taken to gauge 

the cooling effect that the injection of oxygen would have. Oxygen was not 

injected as the injection valve uses an oil lubricant which would burn on 

contact with the oxygen. The conductivity constants of nitrogen and oxygen at 

1 atm and 3OO0K are 0.0260 W/mK and .0262 W/mK and the specific heats are 2080 

J/kgK and 1820 J/kgK respectively. Hence little difference is expected 

between the two gases. The conductivity constant and specific heat for 

hydrogen are .181 W/mK and 14460 J/kgK respectively, so that the value of k/cp 

for both nitrogen and hydrogen is 1.25 x 10 kg/sec m. The values of m y 

for the nitrogen and hydrogen injection are 2.8 x 10 kg/sec and 3.2 x 10 

-5 

00 

-3 -3 

- -1 

kg/sec, so that if K a is of second order to q then the heat transfer 

rates would be expected to coincide. From Figure 45 it is seen that heating 

first occurs at the same distance downstream of the injector, however towards 

the end of the duct the hydrogen injection would appear to give a greater 

cooling rate than the nitrogen injection. This is not understood. 

Conclusions. 

At the higher enthalpies an increase in m y gives greater insulation to the 

thrust surface. At lower enthalpies effects which are second order at the 

higher ar?ttr.=l=ies =re fiew eqally as- i9.porta~t. The shple ~.edel prepcsed fer 

the heating of the thrust surface is qualitatively useful. However, due to 

00 

.- 

the lack of an accurate value of the effective diffusivity constant it is 

limited in its ability to predict quantitative results. 
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SECTION F 

2 STAGE DIVERGENCE. 

Previous experiments with wall injection shc ed low values of thrust 

and specific impulse. This was attributed to two effects.Ref 1. 

Firstly, cornbustion was prevented from spreading to the wall bv 

cooling effects. This leads to a loss of performance which is 

equivalent to hydrogen representing an equivalence ratio of *- 1.5 

being wasted. 

This is a feature of the cold walls in the shock tunnel tests which 

would not be expected to apply in a flight situation. Several means 

are under investigation at the moment to make the shock tunnel tests 

more representative of real conditions, but thev do not applv to the 

present report. They include preheated fuel. heated walls and a 

geometrical configuration using a central injector configured ta look 

like a wall injector in so far as the expansion is concerned. The 

latter Droposal simulates wall injection. out with no heat transfer or 

shear of the fuel iet on the centreline upstream of the thrust 

surface. This configuration is shown in Fig 46. 

Secondly, as the combustion laver is attached to the wall the zone 

over which it generates thrust bv the expansion and jet interaction 

mechanism, Ref 9, is limited. A means of improving this is tc allclw 

extra length in the combustion chamber before expanding the flow. This 

was done in Drevious experiments, Ref 1 Fi% 17. and a limited increase 

in thrust was confirmed. A weight Denaltv would be associated with the 

longer combustion chamber. and it is advantageous to start expanding 

the flow as soon as possible after injection. 
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However, expansion of the jet before complete combustion is achieved 

can have a quenching effect on the flame, and it also slows down the 

spread of the jet across duct. The optimum configur3tion of combustion 

chamber and expansion nozzle is likelv to be a compromise between 

these two effects. 

It was noted in previous experiments that the thrust produced by wall 

injection was not very sensitive to the combustion chamber divergence 

angle when the expansion / jet interaction is absent. It was therefore 

decided to construct a combustion chamber with a 2 stage divergence. 

This is shown schematically in Fig 47. 

The fiFst stage includes a thrust surface inclined at 4 degrees to the 

intake flow. This geometry was shown in Ref 1 Fig 15 to produce 

significant thrust due solelv to compressions from the burning fuel 

jet. Because only a small pressure drop is associated with a 4 degree 

expansion , it was hoped that the jet would continue to mix and react 

and spread across the duct in the first thrust producing stage. CJhen 

subsequentlv expanded bv the second stage the thickness of the jet 

should be such that substantial thrust could be produced by the 

expansion interaction method. 

The advantage over a single expansion of 15 degrees is twofold. 

Firstlv , if the full 15 degrees expansion is introduced too earlv. no 

further combustion will take place. By limiting the initial expansiQn 

to 4 degrees, thrust may be developed in a region that is still 

supporting combustion. 

Secondlv. the final expansion takes place at a point where the iet has 

had time to spread awav from the wall. Combustion creates a region of 

reduced Mach number attached to the wall. and after the passage of the 

corner expansion this region will be at a higher static pressure than 

the free stream flow. This high pressure region is subsequently- 
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eliminated bv a svstem of expansions and compressions. but it creates 

a region of increased wall static pressure downstream of the corner 

which produces a thrust increment. It can be seen from Fig 4e that the 

zone of increased wall static pressure will grow as the jet continues 

to spread away from the wall. It is this principle that give extended 

combustion chambers potential for increased thrust from the 

interaction mechanism. 

Short combustion chamber.4/15 degrees divereence. 

The initial series of tests was done with the model configured as 

shown in Fig 47 with a 25 mm length of constant area duct after 

injection before the start of the 4 degree diverging section. 

In Fig 49 the results of the dual stage divergence are compared to 

previous data for a single thrust surface with 15 degrees divergence. 

It is seen that at an enthalpy of 8.7 Mjikg no improvement in 

performance was produced. 

In Fig 50 the pressure against distance profiles are shown. It is seen 

that despite significant heat release in the 4 degree section. as 

shown bv the pressure rise above fuel off levels. very little net 

thrust is developed on the downstream thrust surface. This compares 

unfavprably Fig 17 Ref I, where an extended combustion chamber 

followed by a single 15 degree expansion gave a large pressure rise 

through the expansion interaction mechanism. 

At an enthalpv of 4.2 Mj/kg no improvement in specific impulse was 

produced. as can be seen from Fig 49 b. However it can be seen frclm 

Fig 50 b that the effects of the reduced divergence are just beginning 

to be felt in the form of combustion induced pressure rises towards 

the downstream end of the first thrust surface. This would suggest 

that somewhat more distance for combustion was required before the 
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start of the 15 degree section. Consequentlv the two stage thrust 

surface was then used with an extended combustion chamber. although 

this does to a certain extent defeat the purpose of the twcl stage 

expansion. which is to obtain thrust in all sections where the fuel is 

burning. 

Long combustion chamber. 4/15 degrees divergence. 

At'an enthalpv of 8.7 Mj/kg no improvement was produced bv the extra 

mixing length. as mav be seen by comparing Figs 5la and 49a for 

specific impulse, and Figs 52a and 50a for P/X dependence. 

At 4.2. Mj/kg a noticeable improvement in performance was gained bv 

adding the combustion chamber extension. This is shown both in the 

increased value of specific impulse . Fig 51b . and also in the 

development of net thrust on the second thrust surface, Fig 52b. 

It would appear that in the higher enthalpv condition .8.7 Mj/kg , the 

temperature after the initial expansion of 4 degrees is sufficiently 

high that combustion is fast , and the heat release is only limited bv 

the quenched zone attached to the wall. The addition of extra 

combustion chamber length does not change this. and no increase in the 

difference between fuel on and fuel off was observed between Figs 50a 

and 52a. 

However at the lower enthalpy condition , 4.2 Mj/kg , it is seen in 

Fig Sob that without the combustion chamber extension sienificant heat 

release only occurs towards the end of the 4 degree section , and not 

much net thrust is produced on either surface. This is thought to be 

due to the longer ignition delay at the lower temperature. In this .. 

case when a longer combustion chamber was used ignition occurred 

upstream of the first expansion , and increased thrust was develap5d 

. :e 

on both surfaces, as seen in Fig 52b 
38 

A similar effect has previously been observed with central injection, 

Ref 9 , where increasing combustion chamber length is onlv beneficial 

at the lower enthalpies. and this mav also be true for wall injection. 

Transverse and parallel injection. 

In an attempt to improve the performance of the wall injected scrami2t 

, 'a modified iniector was constructed with transverse holes as well 

as the parallel iniection throat. It was hoped that the transverse 

mclmentum of the jets would carrv some of the fuel through the boundary 

laver and awav from the quenching effect of the wall. and would also 

increase the mixing rates. 

A schematic of the injector is shown in Fig 53. It consists of a 

series of hciles drilled at 45 degrees to the flow and pointing 

downstream. The percentage of transverselv injected hvdrogen was 

controlled bv changing the throat of the parallel injectclr. and 

setting the hvdrogen reservoir pressure to give the required total 

equivalence ratio. Transverse hvdrogen mass flow rates equal to 27: 

and 79% of that passing through the parallel injectors were produced 

for the 0.9 mm and the 0.1 mm throats respectivelv. No other 

combipations were used. 

Reduction of the size of the quenched zone would be evident in the 

form of increased values of specific impulse. and also comhustion 

would be possible at lower values of equivalence ratio. Combustion in 

a scramjet using fuel frmn a rocim temperature ri.servoir requires 

heating of the fuel to its' ignition temperature by transfer of heat 

from the free stream. 
39 

In the wall injection case this flow of heat is partiallv offset bv 

the flow of heat to the wall. and is also limited by only having one 

fuel air interface. as opposed to two for cen%ral iniectors. II? the 

two dimensional configuration the area for heat transfer frclm the flow 

to the jet is equal to the area for transfer from the jet to the wall. 

and this limits the temperature which may be achieved in the jet. 

However a circular jet propasating transverselv across a duct will 

have a much larger area exposed to the flow . and mav be expected to 

get hotter. 

Transverse injection. long combustion chamber. 

Two stage divereence 

In Fig 54 the results of transverse injection with an extended 

combustion chamber are shown. Thev are reproduced in Fig 55 for the 

4.2 and 8.7 Mj/kg conditions to give a direct comparison with the 

other canfigurations. 

Transverse injection at the 8.7 Mj/kg enthalpy condition again gives 

no improvement over anv of the other results. except at low values of 

equivalence ratio. Combustion with transverse injection occurred at 

lower equivalence ratios than was Dossible with parallel injection 

alone. This is indicated by the two points on Fig 55a at equivalence 

ratios of 0.83 and 1.11. This would suggest that above a certain fuel 

injection pressure the transverse jets are to some extent penetrating 

the boundarv laver. and burning at iower equivalence ratios than is 

possible for wall injection. However. the heat release frcm this 
combustion is not reducing the size of the quenched zone in the 

parallel injected component. and so no increase in specific impulse is 

observed at higher equivalence ratios. 

Also shown on Fig 54a is a data point obtained bv injecting Helium 

instead of hvdrogen. This was done for selected conditions, together 

with the injection of hydrogen into nitrogen test gas. to separate the 

effects of combustion from the physical presence of a jet of foreign 

gas in the flow. In this case a genuine cwntiustion effect appears to 

be present as the hvdrogen is producing significantlv more specific 

impulse than the helium gas. 

At 4.2 Mj/kg combustion at low equivalence ratios is again observed . 

and there also appears to be significant improvement in performance in 

the equivalence ratio range of 1 to 2.5 . as may be seen from Fig 55b. 

Also shown on this figure is are the results of 4.2 Mj/kg tests with a 

transverse component equal to 79% of the parallel injection. the only 

condition for which this was done. This shc?ws no imDr[Dvement over the 

27% case. It had been hoped that a larger proportion of transversely 

injected fuel would lead to more combustion. and higher specific 

impulse. This effect would seem to indicate that the transverse jets 

are not penetrating verv far into the flow, and are still restricted 

by wall quenching. 

Further evidence of this is given in Fig 57. which compares the wall 

pressure and temperature profiles for a single 15 degree expansion at 

an enthalpy of 6.1 Mj/kg for parallel and 27% transverse injection. No 

difference is apparent between the pressure traces , and onlv a slight 

increase in heat transfer with transverse injection was observed. Both 

tests were taken at an equivalence ratio of about 2. where no thrust 

increment was observed with transverse injection. A mcare significant 
41 

result would be at equivalence ratios of order one where the 

transverse injection appears to be effective, but no heat transfer 

data was taken at those conditions. 

It is possible that the transverse jets also require a laver of fuel 

as a thermal buffer to insulate them from the wall, and if this is not 

supplied bv the parallel jet then more of the transverse component 

will be quenched. Another factor which mav be significant is the 

effect of the expansion from the trailing edge of the injector. The 

strength of this expansion is dependent on the amount of fuel injected 

through the parallel throat. Less fuel from the parallel throat would 

lead to a stronger expansion and a region of cooler fuel downstream of 

the injector. 

Lonz combustion chamber 15 degree single expansion. 

A final series of experiments with transverse injection were performed 

with a long combustion chamber and a single 15 degree thrust surface. 

Transverse injection was a the rate of 27% of the parallel component. 

The results are presented in Fig 56. together with the corresponding 

data for the 2 stage expansion with a long combustion chamber. 

Excep.t for the 4.2 Mj/kg case . which only just seems to be on the 

verge of igniting for the single expansion. no difference was observed 

between the two conditions. 
Conclusion. 

In conclusion for this section, the 2 stage expansion can lead to more 

combustion with a short combustion chamber, but this does not lead tcl 

the development of more thrust. presumably indicating that the 

reacting laver has not spread far from the wall. This is shown for a 

low enthalpy condition in Fig 50b and Fig 49b. 

For lclnger ducts the double expansion is able to produce more thrust 

at lower enthalpy than a short one , but the improvement over a single 

expansion with a long duct is marginal. For high enthalpv .8.7 Mj/kg, 

no combination of divergence or combustion chamber length was able to 

produce an improvement with Darallel injection. 

The wall quenching layer appears to be the predominating effect , and 

the effectiveness of staged expansions cannot be properlv measured or 

understood until this problem is alleviated. Several possible 

approaches to this problem are suggested at the beginning of this 

section. 

The use of transverse injection was shown to be partially effective in 

lowering the minimum equivalence ratio for combustion. At higher 

equivalence ratios it did not improve specific impulse.except to a 

slight extent at 4.2 Mj/kg . This was taken to imply that although 

some of the transversely injected fuel may burn at equivalence ratios 

where combustion would not otherwise be Possible, it did nclt reduce 

the amount of parallel injected fuel which was quenched thy the wall. .r 
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No improvement was produced bv injecting a larger proportion of 

transverse fuel. This suggests that the penetration of the transverse 

jets into the freestream is limited. and that combustion still relies 

on an insulating layer of fuel to act as a buffer to the cold walls. 

Because the combustion is again dominated bv wall quenching it is not 

possible to asses the effectiveness of the transverse jets as a mixing 

aid. However heat transfer measurements did not show significant 

change in the heating rates . and it does not appear that it had a 

strong effect. This would again imply that it is difficult to get a 

hydrogen jet to penetrate into the freestream. 

The use of transverse injection into a long combusticln chamber, 

followed by a two stage expansion. did produce the highest values of 

specific impulse vet obtained in the shock tube for wall injection. 

This is shown in Fig 56c where a value of 550 sec was achieved st 

equivalence ratios of -’ 1. 
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SECTION G 

SHOCK STABILIZED COMBUSTION. 

For combustion of fuel in a scramjet three processes are necessary, 

namelv mixing of the fuel and air. heating of the fuel atlove its 

ignition temperature and the allowance of sufficient time fer 

reactions to take place. All these Processes would normallv be 

expected to occur concurrentlv. However, the compression process in a 

scramjet. which consists of an initial compression bv an oblique bow 

shock followed bv further compression at the ccmbustion chamber 

intake; allows for the separation of the first of these processes. 

A schematic diagram of a scramjet is shown in Fig 58. The pressures 

and temperatures after the bow shock will tvpicallv not be sufficient 

to support combustion. If fuel is injected after the bow shc?ck. but 

before the combustion chamber intake. it will not burn, but as the 

intake is likelv to be verv 1ang it mav be expected to mix well with 

the freestream air. It will enter the combustion chamber as a premixed 

fuel air mixture. and some heating of the fuel will have taken place. 

Both fuel and air will be heated bv the combustion chamber intake 

shock waves to form a combustible mixture. Combustion will then onlv 

depend on the ignition characteristics at the temperature and 

pressure after the shock. Ignition bv means of a shock wave is known 

as shock stabilised combustion. 

The combustion chamber can therefore in theorv be made shorter and 

lighter because the mixing and heating stages are completed at the 

intake. An additional benefit might occur if the fuel were to be 

injected from the wall, and could act as a coolant for the und2rside 

of the craft. To prevent choking at the intake a sufficientlv high 
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enthalpv and Mach number flow is required. A preliminarv series of 

tests were performed on the T3 shock tunnel to see if this effect 

could be realised. 

Central injection. Mach 5 flow. 

The purpose of the experiments was to inject hvdrogen into a flow of 

air at conditions which do not sustain combustion. and then to induce 

ignition bv compression through an oblique shock wave. A schematic 

diagram of the experimental model is shown in Fig 58. 

A mach 5 contoured nozzle was used . with which it has previouslv 

been found hydrogen does not burn. A wedge inclined at 5 degrees to 

the intake flow was used to create the shock wave .and reflection of 

the shock from the wall was eliminated bv a 5 degree corner starting 

at the mint where the shock meets the wall. Down stream of the 

expansion fan the flow was contained in a constant area duct formed 

b5tween the wedge and ths sidewall. Equivalence ratios were high.of 

order 5. 

in Fig 59 the results for a range of enthalpies are shown. At the 

higher enthalpies. Fig 59a + 59b. ignition is almost immediate after 

shock reflection. as may be seen bv comparison of the fuel on and off 

pressure profiles. Before the reflection of the shnck frnm the wall no 

pressure increase from combustion was clbserved. 

As the enthalpv is reduced ignition occurs with a significant delav 

after shock reflection, as mav be seen in Fig 59c for an enthalpy of 

3.43 Mj/kg. 

As the enthalpv is reduced further this delay increases. as may be 

seen in Figs 59d and 59e for enthalpies of 2.65 and 2.3 Mj/kg. At 

these last two conditions combustion was unsteady. as determined bv 

the normalised plots of static pressure/stagnation pressure against 
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time. The unsteadvness was in the form of a pressure pulse starting 

down stream and propagating up the duct. These two results should 

therefore be treated with caution. but thev do serve to illustrate the 

effects of enthalpy on ignition delay times. 

As the enthalpv was further reduced to 1.9 Mj/kg no combustion was 

observed at all, Fig 59f. As this was just a preliminarv look at the 

shock stabilised combustion effect, fuel off traces were not taken for 

all enthalpies. As the form of these does not change much with 

enthalpv. fuel off traces from different enthalpies were used where 

necessary for qualitative comparisons. 

The passage of a shock wave through premixed fuel was found therefore 

to be very effective in heating the gas and inducing ignition. Large 

pressure rises were recorded in a flow that would not normallv support 

combustion. An unsteadv combustion effect was observed at the lower 

enthalpies which could be the onset of thermal choking. 
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Wall injection. Mach 5 flow. 

The shock induced combustion concept was also tried with wall 

injection. There were two reasons for this. Firstlv. if this techniaue 

is to be used on a flight vehicle, the fuel will probably be injected 

from the wall in order to provide boundarv laver cooling. 

Secondlv, previous experiments with wall injection in the combustion 

chamber have all been affected bv a cold laver of quenched fuel near 

the wall. It was hoped that by reflecting a shock off the model walls 

this quenched laver might be reduced in size and effect. 

The experiments at Mach 5 were performed for the first reason. 

The configuration of the model was similar to that used with central 

injection. but a shock turning angle of 10 degrees was used. Because 

the shock was reflecting off a thick low Mach number laver of fuel, 

disturbances were apparent upstream of the shock. The model 

configuration for wall injectiQn with shock stabilisation is shown in 

Fig 60. 

Combustion was not possible at this Mach number for wall injection 

without shock compression. The tests were performed at an enthalpv of 

8.7 Mj/kg, and with equivalence ratios from 0 to 5.7. The results are 

presented in Fig 61. and it is seen that it did not burn. The pressure 

rises. upstream of the shock apparent in Fig 61 are attributed to a 

shock wave interaction process rather than to combustion. This is 

confirmed in Fig 62 where the results of helium and hydrogen injection 

are compared at an equivalence ratio of 5.7 for the hydrogen case. 

with the helium being injected at the same pressure. 

The shock ignition process was therefore found to fail for wall 

injection at a condition where it had been most effective for central 

injection, despite being compressed bv means of a stronger shock. This 

.- 
is provisionally being attributed to the quenching effects of the 

walls. This does not preclude the use of shock stabilised combustion 

with wall injection on a flight vehicle with hot walls. 

Wall injection. Mach 3.5. 

At this Mach number the purpose of the experiments was to trv to 

reduce the extent of the wall quenching zone. Before the arrival of 

the shock combustion would already be established in a limited zone 

between the free stream and the cooled laver in the immediate vicinitv 

of the wall. However it is thought that mixing of oxidant penetrates 

beyond-the reacting laver and into the quenched zone. Therefore if the 

tsmperature of this mixed but unburned region could be raised, more 

combustion and heat release might be produced. The quickest and most 

uniform way to heat this region is bv the passage of a shock wave. and 

this is whv shock assisted combustion was tried. 

Because this is a condition where combustion is already possible. the 

effectiveness of the technique would be shown bv a reduction in the 

minimum equivalence ratio at which anv burning takes place. It should 

also result in higher pressure levels in the equivalence ratio range 1 

to 2.5 which is the only region where heat release was found to 

increase with equivalence ratio for wall injectia2n. 

In Fig 63 the pressure profiles are shown for an enthalpy of 8.7 Mj/kg 

at an equivalence ratio of 0.91. This is a condition which d~es not 

burn with a constant area duct because the equivalence ratio is too 

high. It still does not burn with the shock heating. showing thst 

there is not significant reduction in the size of the quenching zone 

at this condition. 
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At enthalpies of 6.1 and Mj/kg more comprehensive data was obtained. 

These results are displaved in Fig 64 and Fig 65. Again it is seen 

that down stream of the shock, in the constant area duct sectiGn. 

there is no improvement in performance over the results quoted in Ref 

1. This applies both to the pressure rises and the minimum quivalence 

ratio at which it will burn. 

A possible explanation for the different results the shock wave has on 

central and wall injection mav lie in the effects the phvsical size 

reduction of the compression has on the heat transfer rates in the two 

configurations. 

In the central jet heat is flowing into the mixing layer from both 

sides, and after the passage of the shock wave the driving temperature 

difference is increased, and the distance the heat has to be 

transferred over is reduced. There is therefore a significant increase 

in %he rate of heat transfer to the mixing laver which, when coupled 

with the corresponding pressure increase, initiates combustion. 

For wall injection the temperature of the jet is determined by a 

balance of the heat flow from the freestream to the mixing laver .and 

€rom the mixing layer to the wall. This was mentioned in Section F 

when discussing the effect of transverse injection. 

Zhilst the passage of the shock wave wiii increase the flow of heat 

from the freestream, it will also for the same reasons increase the 

heat transfer from the jet to the wall. The temperature rise in a wall 

injected jet will therefore be smaller than that produced in a central 

iet. From the experimental results it would appear then that the net 

result of the compression is that the temperature is not much 
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increased in those regions where the fuel and air have mixed to 

combustible proportions. but are prevented from burning by the cooling 

effects of the wall. 
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TABLE OF TEST CONDITIONS 

H M T P 

MJ/kg K kPa 

8.70 

8.70 

8.70 

6.10 

6.10 

6.10 

4.20 

4.20 

4.20 

3.43 

3.43 

2.73 

2.65 

2.65 

2.65 

1.94 

1.90 

1-90 

1-19 

1.07 

1.03 

3.50 

4.50 

5 

3.50 

4.50 

5 

3.50 

4.50 

5 

4.50 

5 

3.30 

3.50 

4.50 

5 

3.30 

4.50 

5 

3.40 

3.40 

3.40 

2500 

1740 

1500 

1700 

1200 

1000 

1100 

750 

650 

625 

540 

970 

700 

480 

410 

705 

400 

350 

445 

415 

400 

160 

32 

20 

160 

32 

20 

160 

32 

20 

30 

20 

10 

120 

30 

20 

15 

30 

20 

24 

24 

24 
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REFERENCES. 

1. R.G.Morgan. A.Paul1, N.A.Morris. R.J.Stalker. 

"Scramjet sidewall burning - preliminarv shock tunnel results." 

University of Queensland. 

Department of Mechanical Engineering Research Report No: 12/85. 

2. 'M.P.Netterfield, R.J.Stalker, R.G.Morgan. 

"The reflection of an expansion fan from a Mach mumber gradient in a 

scramjet exhaust. " 

University of Queensland. 

Department of Mechanical Engineering Research Report No: 7/84. 

3. W.D.Haves. R.F.Probstein. 

"Hvpersonic flow theorv. 'I 

Academic Press. New 'fork and London. 1959. 

4. J.L.Stollerv. G.T.Coleman. 

"A correlation between pressure and heat transfer distributions at 

supersonic and hypersonic speeds. VI 

Aeronautical Quarterlv, Vol. 26. pp 304-315, 1975. 

5. J.A.Lordi. R.E.Mates, J.H.Moselle. 

NASA rep. NASA CR-472. 1966. 

6. R.A.East, R.J.Stalker. J.P.Baird. 

"Laminar flet plate heat transfer measurements from a dissociated high 

enthalpv hvpersonic air flow". 

Department of Aeronautics and Astronautics, 

.- 
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Southampton Report No. 338. 1977. 

7. A.Paul1,N.A.Morris.R.G.Morgan.R.J.Stalker. 

"High reynolds number heat transfer to the cold walls of a model 

scram jet". 

To be presented to the 9th Australasian Fluid Mechanics Conference. 

Auckland. November 1986. 

8. J.S.Evans. C.J.Schexnavder Jr.. H.L.Beach Jr.. 

"Application of a two - dimensional parabolic computer program to 

prediction of turbulent reacting flows." 

NASA Tech. Paper 1169, March 1978. 

9. R.J.Stalker.R.G.Morgan. 

"Supersonic Combustion with a Short Thrust Nozzle". 

Journal of Combustion and Flame,Vol. 57. No. 1. 

JulV 1584, pp 55-70. 

10. Allen G. McLain and C. S. R. Rao 

"A Hvbrid computer Program for Rapidlv Solving Flowing or Static 

Chemical Kinetic Problems Involving Many Chemical Species." 

NASA Technical Memorandum X-3403. Julv 1976. 

11. Casimir J. Jachimowski and Allen G. McLain 

"A Chemical Kinetic Mechanism for the Ignition of 

Silane/Hydrogen Mixtures" 

NASA Technical Paper 2129. February 1983. 
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ACKNOWLEDGMENTS. 

This work was performed under a grant from the NASA Langlev Research 

Center. Hvpersonics Propulsion Branch. 

The experimental work was performed in the shock tunnel of the Phvsics 

department of the Australian National Universitv. and the assistance 

of'many members of the staff for all the back up suppclrt required for 

the operation of a maior test facilitv is greatlv apprexiated. 
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55. 

APPENDIX 

Additional work has been carried out on the simulation of the 

combustion of silane/hvdrogen mixtures in a constant area duct. 

Using a one dimensional chemical kinetics program (Ref. 10) and 

the reaction scheme postulated bv Jachimowski (Ref 11.). results 

have been compiled for four intake temperatures at an intake 

pressure of 2OkPa and a mach number of five. A 20% silane in 

hydrogen fuel was considered for these examples. Of particular 

interest is the effect of freestream oxygen radical concentration 

on the combustion process. Comparisons are made with experiments 

involving central injection into a constant area duct. 

Because the program is one dimensional and applies to premixed 

gases onlv. its use is restricted to the examination of the 

effect of chemical kinetics on ignition delav times rather than 

details of pressure profiles when comparing with experimental 

results . 

Figure 67a shows that at 1500K or a freestream enthalpy of 8.7 

MJ/kg, increasing the radical oxvgen has little or no effect on 

the ignition delay time. Results from this simulation compare 

favourably with the experimental results for a 20% mixture as 

shown in figure 11, given that injection takes place at 

X=W.5mm. (ALPHA is the percentage mass of freestream oxygen that 

has dissociated). It is suggested therefore that at this high 

enthalpv. the high dissociation expected in a shock tunnel 

(ALPHA=10%) has no noticeable effect on the ignition delav. 
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56. 

~t iaGGi( or a freescream entnaipv of 6.1 ~J/kg. the shock tunnel 

produces flows with an ALPHA of around 1.5%. According to the 

simulation (figure 67b1, this would bring the rapid pressure rise 

to around 20 cm closer to the point of injection than when 

compared to a flow with no oxvgen dissociation. Figure 12 

shows a high pressure is measured closer to the point of 

injection than the 20 cm predicted for a zero dissociation flow 

which suggests that free radical oxvgen does in fact aid 

combustion in this case. 

Figure 67c shows that at 650K (4.2MJ/kg), only small amounts of 

free radical oxvgen have a very significant effect on the 

ignition delay time predicted bv the simulation. At this 

enthalpv. non-equilibrium calculations have shown that an ALPHA 

(2f 0.18% is expected in the shock tunnel. It can be seen bv 

comparing the results for a 20% silane mixture shown in figure 13 

with simulation for ALPHA=0.2% in figure 67c. that ignition delay 

times are around the same. 

At a lower temperature of 540K. ignition delay times are 

dramatically increased for the lower radical concentrations. as 

predicted bv the computer simulation (figure 67d). However. 

figure 14 shows that in experiments with an ALPHA of 0.13%. the 

20% silane mixture ignites at about 1.5 cm from the point of 

injection. This compares with about 65 cm for premixed gases as 

predicted by the 1-D program. Similarly, at the lower temperature 

of 410K (figure 15), it is seen that ignition takes place at 

about 1.5 cm froin the injection point, vet the oxygen 

dissociation is calculated to be onlv @.06%. 
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57. 

if +La Experiments suggest th,eref=re that L11c rzacti~ii ~~ih~rn~ is 

valid, then freestream radical production at high enthalpies 

(1500K freestream temperature) has no appreciable effects on the 

ignition of 20% silane/hvdrogen mixtures. 

At. the intermediate freestream temperature of 650K, oxvgen 

radical concentrations play an important part in the combustion 

process of silane/hydrogen mixtures therefore shock tunnel 

simulations may not be representative of real flight conditions. 

At the lower temperatures of 540K and 41@K, experiments showed 

short ignition delav times despite the verv low free radical 

oxvgen concentrations produced by the shock tunnel. This suggests 

that combustion is occurring because of other effects. These mav 

be the high temperature produced by the boundary laver on the 

injection strut and/or production of free radical oxvgen in this 

boundary layer. If this is the major ignition mechanism. then 

freestream radical oxvgen produced bv shock tunnel flows will not 

have a significant effect on the combustion process. Rather, 

ignition will be determined tv the presence of a central 

injection strut in hypersonic flows. 
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$4 

0 

U 

-4 

C 

0 

E 

0) 

Ll 

3 

v) 

II) 

al 

Ll 

a 

0 

-4 

U 

m 

U 

v) 

0) 

Y 

cp 

u 

C 

H h N 0 C \ 

v-! 
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1 

: 

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Q) 

U 

rd 

w 

U 

3 

tn 

a 

Q) 

Y 

C 

a, 

5 

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P/Ps 

0.10 I I I 4 I I 1 1 I 1 r F 0 ,0 

.O. 

100 2 00 3 00 400 500 

X,mm 

Ofuel off, T= 400K 

020% SiH,/H, injection, TI = 400K 

020% SiH,/H, injection, TI = 415K 

A 20% SiH,!H, injection, TI = 445K 

Fig. 3. Ignition limit of 20% SiH,/H, mixtures at mach 3.4, @ - 0.3, 

PI = 24kPa 

Time, 307ps per division 

Stat ion X,mm 

(3) 130.5 

(4) 155.5 

(5) 180.0 

(6) 205.5 

Fig. 4. Illustration of thermal choking at low Temperatures, 20% SiH,/H,, 

T = 415K, P = 24kPa, @ 0.3. - 

I I 
100 200 3 00 400 500 

X,mm 

I - I 

0 Fuel off, TI = 705K 

0 20% SiH,/H, injection, TI= 705K 

Fig. 5. Ignition limit of 20% SiH,/H, mixture at mach 3.3, P = 15kPa, I 

0 = 0.78 

PIPS 0.1 t 

I- 0 

0 100 

-- 

200 3 00 400 500 

X,mm 

OFuel off, T = 970K 

020% SiH,/H, injection, T = 970K 

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Fig. 6. Combustion of 20% SiH,/H, mixture at mach 3.3, P = lOkPa, # = 1.58 

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0.010 

P/Ps 0.005 

0' 

I 1 I 4 1 I 1 I I 

- 

- - L 0 - 

- 0. 0. 

Om 

*a 

0 

- 

a o,08g - O. - 

- 

- F3n 0 OB 

k A. ao 

I a 

c - 

- - I 1 I I 1 1 1 I 1 L 

Fig. 7. Effect of concentration of silane on Combustion at mach 4.5, 

Hs = 3.43MJ/kg, P = 30Wa, T 6.25K I I I= 

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l 

I 

P, kPa 

150 

1 I I I I I l I I 

c 

75 

1 I 1 1 I I I I 1 I 

Time, 307ps per division 

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Stat ion - X,mm 

3 190.5 

4 233.5 

5- 258.0 

6 284.0 

Fig. 8. Constant pressure burning of a 10% Silane Hydrogen mixture at mach 4.5, 

Hs = 2.65MJ/kg, P, = 30kPa, T = 480K I I I 
PIPS 

0.010 I I I b I I 1 1 4 

c *I 0.005 

0. 0 a 

.. 0 

0 

.AAA 

0 

0 

0 OA A 

01 1 1 I I 1 1 I 1 I 1 

140 280 420 660 7 00 

X,mm 

0 Fuel off 

0 5% SiH,/H, injection, # = 0.57 

0 10% SiH,/H, injection, # = 0.65 

A 20% SiH,/H, injection, # = 0.58 

Fig. 9. Effect of concentration of silane on combustion at mach 4.5, 

Hs = 2.55MJ/kg, P = 30kPa, T = 480K 

I I 

0 Fuel off 

0 5% SiH,/H, injection,-.# = 0.72 

10% SiH,/H, injection, 8 = 0.63 

A 20% SiH,/H, injection, # = 1.07 

Fig. 10. Effect of concentration of silane on cornbustion at mach 4.5, 

Hs = 1.90MJ/kg, P - 30kPa, T = 400K 

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I P9kPa 

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.' 

,0050 1 I I 1 I I I 1 I I i 

0 I I I I I I 1 I 1 

3 00 450 600 7 50 0 150 

X,mm 

0 fuel off 

0 5% SiH,/H, injection, Q = 1.3 

A 20% SiH,/H, injection Q = 1.21 

Fig. 11. Effect of concentration of silane on combustion at Mach 5, 

H = 8.7 MJ/kg, PI = 20 kPa, TI = 1500 K S 

1 1 I I I 1 I I 1 

150 450 600 7 50 3 00 

X,mm 

0 fuel off 

2.5% SiH,/H, injection, ,$ = 1.08 

A 20% SiH,/H, injection, ,$ = 0.90 

Fig. 12. Effect of concentration of silane on combustion at Mach 5, 

H = 6.1 MJ/kg, PI = 20 kPa, TI = 1000 K 

S 

(Note: 

pressure information) 

Pressure is absolute due to lack of stagnation 
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PIPs 

0 fuel off 

2.5% SiH,/H, injection, @ = 1.03 

0 5% SiH, H, injection, @ = 0.92 

0 10% SiH, H, injection, I$ = 0.87 

A 20% SiH, H, injection, @ = 0.99 

Fig. 13. Effect of concentration of silane on combustion ar' Mach 5, 

H = 4-2 MJ/kg , P = 20 kPa, TI = 650 K 

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S I 

L 0 -r 

- 

- .I 

0.00375 - 

S 

- - 

- - 

- - 

0. 

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I 1 I 1 I 1 I I 

- 

1 

3 00 450 600 750 0 150 

x,mm 

0 fuel off 

2.5% SiHt,/H, injection, $ =0.95 

0 5% SiH,/F, injection, $ = 0.92 

0 10% SiH,/H, injection, I$ ~0.77 

A 20% SiH, H, injection, $ = 0.87 

Fig. 14. Effect of concentration of silane on combustion at Mach 5, 

Hs = 3.43MJ/kg, P - 20 kPa, TI =540 K 1- 
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- - 

- 

I I I I I I I 

- 

.. 

.005 1 I I I I 1 1 1 I 1 

.' 1 .0025 

t 00 

x ,mm 

0 10% SiH,/H, injection, 4 = 0.91 

A 20% SiH, H, injection, 4 = 0.68 

Fig. 15. Effect of concentration of silane on combustion at Mach 5, 

H = 2.65 MJ/kg, PI = 20 kPa, TI = 410 K 

S 

0.005 I I I I I 1 I 1 I 1 

i 0.002 

A 20% SiH,/ H, injection, $I = 0.83 

Fig. 16. Combustion limit of 20% SiH,/H, at Mach 5, H = 1.90 MJ/kg, 

TI = 350 K S 
P/Ps 

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.0050 

.0025 

0.0000 

.0050 

.0025 

I I I I I 1 I I L 

I 

- c 

- 

- - 

- - c O$ 

- (3 

€3 0% 8 

- b 0 on - 

- B f3 - 

1 I I 1 I I I 1 1 

x,mm 

0 

n fuel off, H = 4.2 MJ/kg 

0 

S 

17.4% Ar/H, injection, Hs = 4.2 MJ/kg 

Fig. 17. Central injection of an a.rgon hydrogen mixture at Mach 5, 

TI = 650 K, PI 20 kPa, @ Z 1. 

Fig. 18. Central injection of an argon hydrogen mixture at Mach 5, 

TI = 1500 K, P = 20 kPa, @ z 1 I 
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nozzle extension 

injector 

edge of expdnsivn ian ar: mach 3.5 

datum 

\\\\\\M 

r 

/ 

r 

I 

Y 

\ 

113.5 

Fig. 19. 15 degree divergence thrust surface without cowl 

datum 

'A /I/////////// ///,/ 

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I 

/ I 

/ I 

YV 

/ 

/ 

Fig. 20. 15 degree divergence thrust surface with cowl. 

Yt ' 

I 

Fig. 21. 15 degree divergence symmetrical thrust surfaces 

Scale 1:4 
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Mach Number 0 compress ions 

-------- expans ions 

expansion wave 

Compression 

waves 

central expans ion 

inject ion waves 

strut 

Fig. 22. Thrust contribution due to an expansion wave passing through a Mach 

number gradient 

injection 

11 strut fuel - jet 

//// f/ 

h 

waves 1 

Fig. 23. Thrust contribution due to a burning fuel jet 
PIPs 

I 

0.2r 

0.1 

0. 

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1 I I 1 I I I 

L 

I I 

- 

- 

- I- E 1- r o r 

B 

L 

- - 

- 

- A R -. 

b 0 

A 

- 

1 1 I I I I I 1 1 

P/Ps I 

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1 

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I 

0.2 I I I b 8 I 1 1 1 

0.1 

0 

t I 

0 50 

y,mm 

100 

t] fuel off 

A H, injection, $ = 1 

Fig. 24. Pitot profiles across an open duct with 15 degrees divergence, 

Hs = 6.1 MJ/kg, MI = 3.5, T = 1700 K, P 

I I = kPa 

Y,m 

o fuel off 

H2 injection, $ = 1.27 

Fig. 25. Pitot profiles across an open duct with 15 degress divergence, 

Hs = 8.7 MJ/kg, M = 3.5, TI = 2500 K,P - 160 kPa 

I 1- 
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PIPs 

S PIP 

0.1 

0.05 

0 

0 fuel off 

A H, injection, + = 1.0 

0 H injection, + = 1.85 

Fig. 26. Pitot profiles across a closed duct with 15 degrees with 

divergence, H = 8.7 MJ/kg, M = 3.5, TI = 2500 K, P = 160 kPa S I 

I 

0.1 

0.05 

0 

Fig. 27. Pitot profiles across a symmetrical duct, 15 degress 

divergence, Hs = 8.7 MJfkg, MI = 3.5, TI = 2500 K, 

PI = 160 kPa 
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kd.9 I 

30' 

rA. Standard nozzle B. Source flow nozzle 

Fig. 28. Central Injectors 

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A. Standard nozzle B. Source flow nozzle 

Fig. 29. Wall injectors 

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3t 

1 PIP, 

I 

I 

Fig. 30a. Pressure profiles for central injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

I 

I I I I I I I I I -d 

- + cylindrical nozzle 

- 

I 

- 5s - source-flow nozzle 

- 

L H = 4.2 MJ/@ 

I= 

- -I. 

+ 

-s a t I 

- - Cp = 1.38 

- I 

M = 3.5 -. 

- 

b I Ill11 I II II I 

200 II 

I 

I 

q 

W/cm2 

cylindrical nozzle into a constant area duct. 

* II 111111 IIII IIIIII I 

- - + cylindrical nozzle 

- - source-flow nozzle 

- - 

- - 

H = 4.2 MJ/kg - ++ -s 

- *+ Ir, - 

- - = 1.38 9 .8 '3- '2 

I' M = 3.5 

- sr - 

- - v 1 I I 

+ I I I I I 
I 

1 

I 

I 

I 

I 

I 

JW/cm' 

Fig. 31a. Pressure profiles for central injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

I/ F 

+ cylindrical nozzle j 

source-flow nozzle 

H = 8.7 MJ/kg S i Q = 1.72 

+ + .$ 

.+ .++: ..I- .+. 

t+ 

+m 

-?-. .++ 

I 4- E 

3 = 3*5 

m 

I' 

L. I I I I 1 I I I 1 

0 x mm 500 

Fig. 31b. Heat transfer rates for central injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

+ cylindrical nozzle I 

8 source-f low nozzle 

H_ = 4.2 MJ/kg 

4 = 1.48 

N 

+I ++ m 

I 1 

+ 

mm am 

= m+ 3 i-im 

I t 

M = 3.5 1 

I I I1111 I I1 I1 I 

I 

0 x mm 500 

Fig. 32a. Pressure profiles for wall injection from a source-flow and a 

cylindrical nozzle into a constant area duct. I 
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W/cm2 

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B 

I 

8 

I 

- - +r source-flow nozzle 

,A+ 

+ .++ S 

m 

I I= 3 - 

-. - H = 4.2 MJ/kg 

=+ P 

- '+ + - 

@ = 1.48 

- == + - I 

-. +++I - M = 3.5 

- '+ 

- 

I c I I I I I A 

PIP, 

Fig. 32b. Heat transfer rates for wall injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

- + cylindrical nozzle 

- 

- source-flow nozzle 

- H = 8.7 MJ/kg 

i ~-- --~--~--.~--.I-r- --- --~--- 2- 

I 

- S 

- 0 Q = 2.0 

- - 

M = 3.5 

j I "m +v 

I+ 4- + mu 

I =++ ? 

+ 

- - 

- - 

I 11111 1 I1 I1 I 

Fig. 33a. Pressure profiles for wall injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

100 

4 

W/cm2 

-b r* 

=+ 'A 

.. 

+ 

S 

H 

@ 

M 

cylindrical 

source-flow 

= 8.7 MJ/kg 

= 2.0 

= 3.5 

0 500 

Fig. 33b. Heat transfer rates for wall injection from a source-flow and a 

cylindrical nozzle into a constant area duct. 

nozz I I 
200 

I; 

I + 

+ 

++ 

f + + + + 

i 

m mm 

mm mm 

m mm 

I= I= 

+ laminar prediction 

@ measurcd 

H = 8.7 MJ/kg S 

M = 3.5 

Fig. 34a. Predicted and measured values of the heat transfer rate for a 

constant area duct with injector. 

+ 

1 -I---.- -..--.--- - .--. - 

++ j + laminar prediction 

loo r I r 

+ 

I 

I 

+ + 

m m 

+ + + 

+ 

+ i 

\ 

I= 

t 

i 

I 

measured 

H = 6.7 MJ/kg 

M = 3.5 

S 

Fig. 34b. Predicted and measured values of the heat transfer rate for a 

15' diverging duct with injector. I 

I 

I 

P' 

1 

I 

100 

4 

cm2 

0 

+ 

S 

H 

M 

laminar prediction 

measured 

= 4.2 MJ/kg 

= 3.5 

Fig. 34c. Predicted and measured values of the heat transfer rate for a 

15' diverging duct with injector. 
4- 

W/cm2 

I 

I 

I 

I 

I 

I 

I 

I 

- 

- 8 @ = 2.49 

- - 

- H = 4.2 MJ/kg I - S 

- I+ ‘I - 1.6 mm throat injector 

- ++t +

I 

- ++ M = 3.5 

mI - 

... m +*I- - 

I I mr *++ I+ I I I I i 

- +-t’ - 

W/cm2 

1 oo ---___.-. ~~, T- , TT T7 7-,7T- ” --- -- - 

- 1- - 

+ a--+ +: - 

F 

4 

- m.H - 

- I - 

I =? =* =I 

++ +++ 

- I - 

- - + I ++ 

- I 

- +++ - 1 I I I 1 1 I I I 0’ 

+ @ = 1.39 

@ = 0.81 

H = 8.7 MJ/kg S 

0.3 mm throat injector 

M = 3.5 

I 

I 

1 

1 

W/cm2 

4- 6 = 4.6 

@ = 3.45 

H S = 8.7 MJ/kg 

1.6 mm throat injector 

M = 3.5 

x mm 500 

Fig. 37. Heat transfer rates for the injection of hydrogen into a nitrogen 

test gas in a constant area duct. 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

4 

W/cm2 

- -4 

+ +t++ - = 4.75~10 kg/sec 

mo yo - I *q 2 + - 

- I - H = 8.7 MJ/kg 

- I + += S 

- 

- +&+" M = 3.5 + + - I 

I Ram - 

I i 

- *+ I I I 1 I 1 I I I 

Fig. 38. Heat transfer rates in a constant areaduct when hydrogen is 

injected into a nitrogen test gas with different injectors. 

+ 0.3 mm throat 1 0.9 mm throat 

100 

4 

W/cm2 

-- TT----TTTTTT-T'T-IIIITT-r 

- -3 

- 4- h0 yo = 1.3~10 kg/sec 

- 8 io Yo 

- - 

- = 2.5~10 -3 kglsec 

- - 

- H = 4.2 MJ/kg - S 

- 

- M = 3.5 it+,' * *= + p' - + Itf+ p 

+ - 

- *++I IC+ - 

- 

I 

- I II I I 1 I I I A 

+ 0.9 mm throat 4 1.6 mm throat 

mixed fuel I 

q=q 0 I .- 

T=dT-y+T 

YO 

W 

I unmixed fuel 

i X 

W 

T=T 

Fig. 40. Schematic of simple heat transfer model. 
a yo 

kglsec 

5 

2.5~10 

AT 

YO 

K /m 

8.7 M3Ikg 

.- /' 

..... I 

~ ....... .- .... -- ....... ................ ..... __-- .. -- -. 7 I 

...... 4.2 MJIkg I 

I 

8.7 M3/kg 

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

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

,. 

i \ o:T''.i 

*. .. -. .... 

........ .................... ....... ................ 0.9 I ............. 

~ 

..a. 

..e. 

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

I 

.*e 1.6 / ..... 

Fig. 42 The variation of - AT with A. 

YO 

M=3.5 
PIP, I 

I 

I 

0' 

Fig. 43a. Pressure profile for the injection of hydrogen into a constant area duct I and a test gas of air and nitrogen 

.... 

- 

- U Air test gas - I - I 

- m - H = 8.7 MJ/kg 

- S I + 

+I + 

+ - !! I$ = 1.35 + 

- I - - I - M = 3.5 

- - I II I I I 1 I 

----rrr1m-mm---- ..- 

- 

4 *% + - 

-I- .+ . - 

- 

- 

m+ m+t 

+!m m. - + += . ++ 

I 

I 

- A- + I Am* + 

- 

- &+ I 

- I mI 0- I I 1 1 I I 

~~ 

Fig. 43b. Heat transfer rates for the injection of hydrogen into a constant duct I and a test gas of air and nitrogen 

+ N, test gas 

Air test gas 

H = 8.7 MJ/kg 

I$ = 1.35 

M = 3.5 

I 

I 

I 

I 

I 

I 

2- 

PIP, 

I I I I I I i I I A 

- - -k N, test gas 

- - 

- mm I - Air test gas -. mm 

- 

H = 8.7 MJ/kg - S m+ 

mm ++ + - + 

- +d - $ = 2.6 + 

++ + -. - M = 3.5 

- - 

- - i I I1111 I II II I 

Fig. 44a. Pressure profile for the injection of hydrogen into a constant area duct 

and a test gas of air and nitrogen 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

+ &%" - 

71-n-n-n--17171T1--- 

+ N, test gas 

Air test gas 

I 

I 

H = 8.7 MJIkg 

+ = 2.6 

.+ M = 3.5 

I S 

+ * ++ + 

++ + 

I 

I 

0 x mm 500 

Fig. 44b. Heat transfer rates for the injection of hydrogen into a constant 

area duct and a test gas of air and nitrogen. 

7 1 F"TI-lT1 r-rr'TT-i..r -,-. -,--, 

loo r--- 

W1cm2 I L 

+ -t. -I 

m 

I 

+ rho 

m lilo 

H= 

S 

M= 

= 2.8~10 -3 

YO 

-3 Yo = 3.2x10 

8.7 MJ/kg 

3.5 

kglsec 

kglsec 

Fig. 45. Heat transfer rates for the injection of nitrogen into air and hydrogen 

into nitrogen at similar values of +I 

+ nitrogen - air 

0 yo - 

hydogen - nitrogen 
I Shock tube nozzle extension 

I 

I 

I 

1 

I 

I 

I 

I 

I 

I 

P,.-.L..-&: -- I 

UVlIlUUDLlVil 

/p r o du c t s 

r ’ expanded adjacent I -- 

to thrust surfaci 

Fig. 46. Proposal for central injector expanded from’the centreline 

in order to simulate wall injection without quenching 

COWL 

INTAKE 

____, 0 

11’4 J fi L 1 UK Fuel layer 

Fig. 47. Dual stage expansion I 

corner expansion 

Free streem - high Mach Number 

+ 

Reflected expansion ’’ / FUEL 

’wa 1 1 

Fuel on 

1 Fuel off -f ------ 

I 1 I X -i ‘ Region of net thrust production 

Fig. 48. Corner expansion/wall jet interaction 
sp (sec) - I I 

I 480 1 + 2 stage 4/15' 

0 1 stage 15' 

't I 320 } p 

t 

0 4 8 12 16 

400[ + 2 stage 4/15' 

300 1 0 1 stage 15' 

200 I 

100 

@' 

b. H = 4.2 MJ/kg S a. H = 8.7 MJ/kg S 

Fig. 49. Isr, /@ Comparison of dual and single stage expansions. 

I 

I Parallel wall injection, short combustion chamber, M = 3.5 I 

I 

Strikes indicate start of diverging sections 

I( PIPs I .025 [ 1 I !

p @ = 2.18 

P @ = 1.6 

o @ =-Q 

x,mm 

0 $I = 1.82 

0 $=O 

+ Helium injection 

(mass flow rate of heliun 

= 1.29 x mass flow rate of I at @ 1.82) H2 

0 700 x,mm 

B- Hs = 4.2 MJ/kg I Fig. 50. PIX. Dual stage expansion, short combustion chamber. 

Parallel wall injection. MI = 3.5 
I 

I 

I 

I 

I 

I 

A. Hs = 8.7 MJ/kg B. H = 4.2 MJ/kg 

Fig. 51. I IQ. Dual stage expansion. Long and short combustion c 

S 

'P Parallel wall injection. M = 3.5. I 

I 

I 

1 

I 

Strikes indicate start of diverging sections 

0 4 = 1.82 

+ Q = 1.18 

0 4=0 

+ 

+. 9 

98'0 Q +# 

4 

9 

.0125 

L I 

L---.L- I Ill1 I1 If IIIII I 

0 700 x,mm 

B. H = 4.2 MJ/kg S 

I 

1 Fig. 52. PIX. Dual stage expansion with long combustion chamber. 

Parallel wall injection. MI = 3.5. 

hambe rs. 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

Transverse jets 

INTAKE FLOW 

Model wall 

Fig. 53. Arrangement for mixed parallel and transverse wall injection 
I 

I 

I 

I 

I 

I 

1 

I 

- 

- 

- ~~~~ ,-,11 

100 - 

- 

-1 I I t I 

He 

Note: 27% transverse means 

not of total mass injected 

that transverse component is 

27% of parallel component, 

+ + 

+ 

A. 8.7 MJ/kg 27% transverse 

Isr, sec 

+ 

+ 

0 1 2 3 4 5 

B. 6.1 MJ/kg 27% transverse 

sp sec 

1000 I'--I-1-.-77--..7 

I 

I I 

500 1 

I 

+ 

o+ + 

0 

0 

- 0 I I I I 

1 2 3 4 0 -I 

(2. 4.2 M/ks 

+ H, , 27% transverse 

H, , 79% transverse 

Fig. 54. I 7~. Transverse wall injection dual stage expansion. 

SP 

Long duct. MI = 3.5 
!I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

SP 

sec 

1000 

500 

0 1 2 3 4 

B 4.2 MJ/kg 

0 2 stage expansion, long duct, 79% tranverse 

+ 2 stage expansion, short duct, 27% transverse 

0 2 stage expansion, long duct, parallel 

0 2 stage expansion, short duct, pararrel 

1 stage expansion, short duct, parallel 

500 

400 

300 

2 00 

100 

Fig. 

T 

A 8-7 MJ/kg 

55. I /$. Various configurations of wall injection M = 3.5 SP I 
T 

I 

I 

I 

I 

I 

SP 

500 I I t I I I I I 1 I 

o* 0 

0 0 

0 

A 0.7 MJ/kg 

1000 

500 

I 

I Fig. 

. 

1 2 3 4 54 

B. 6.1 MJ/kg 

I 

'P sec 

1000 

500 

I 

I 

+ 

** 

-, .- 

1 2 3 4 54 

C. 4.2 MJ/kg 

56. ISPI$. Long duct 2 + 1 stage expansions 

27% transverse MI = 3.5 

i3 

L 

15' single expansion 

Dual 4/15' expansion 
Start of expansion I 250. I I I ,J I I 1 I kPa 

I .I 

- 

- 

3 

++ I 0 

0 + 

0 

A. PAX 

I mm 

600 X 

mm 

600 X 

I$ = 2.06 27% transverse l 

injection ~ 

+ I$ = 2.04 parallel , 

I 

inject ion I 

B. TIX 

Fig. 57. Comparison of parallel and transverse wall injection. 

Long duct, single 15' expansion. 

I 

Hs = 6.1 Mj/kg , M = 3.5 

I 

I 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

,I 

Int 

Injector 

J/ 

- Shock wedee " 

/ -- e- 

.ake M5 Mixing Constant area 

section combustion section 

Fig. 58. Experimental shock induced ignition ri-g 

for central injection 
1 

I 

I 

I 

I 

I 

1 

I 

I 

I 

I 

I 

I 

I 

1 

I 

1 

I 

1 

I 

100 

50 

100 

50 

0 7 50 1 Xmm 

m@=O 

+ @ = 6.1 

A. H = 8.7 Mdkg S 

.01 

.005 

mm 

0 750' X 

.@=O 

+ l$ = 3.9 

C. H = 3.43 MJ/kg S - 

PIPs hock reflection 

.Ol I1 IIllli 

+tt t 

t 

.005 1 + 

t 

t 

IO 

m @=O 

+ @ = 3.8 

E. Hs = 2.3 MJ/kg, fuel on 

P 

KP a 

mm 

0 750 X 

0 (H = 8.7 MJ/kg) m@= S 

+ I$ = 4.8 

B. H 6.1 MJ/kg , fuel on 

S 

mm 

0 750 k 

a @ := 0 (8.7 MJ/kg) 

+@=4 

D. H = 2.65 MJ/kg , fuel on 

S 

XI 

mm 

+ l$ = 3.7 

F* Hs = 1.9 MJ/kg 

mm 

750 X 

Fig. 59. PIX. Shock stabalised combustion, central injection 

M = 5, 5 turning angle shock. T 
I 

I 

I 

I 

I 

I 

1 

I 

I 

I 

1 

I 

I 

I 

I 

I 

I 

Shock inducing wedge 

I 

\\\ I I 

AIR FLOW 

Injector 

Fuel layer 

I 

INSTRUMENTED 

SURFACE 

Fig. 60. Wall injection shock induced ignition rig. 

10' turning angle 
1 

I 

1 

I 

I 

I 

I 

1 

1 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

.01 

005 

I I I I I I I 

- 

- 0 Q=O 

x Q = 1.95 

0 Q = 3.34 

+ Q = 5.7 

- 

- 

- 

I 

+++ + 

- a* 

- ++ +* Q!l e:+ - 

00 

- *e:* x x - 

4 

- + eau P - 

P X 

U - - 

h I I 1111 I I II Ill I I I 

.01 

.005 

0.1 

n:05 

Fig. 61 P/X shock stabilised wall injection. 10" shock. Hs = 8.7 MJ/kg, M I = 5 

PIP 

+ H2 

CJ He 

x,mm 

injection = 5.7 

injection at the 

same pressure 

Fig. 62. PIX shock stabilised wall injection. 10' shock. M = 5, 

Hs = 8.7 MJ/kg. Comparison at H and H2 injection. I 

e 

0 750 k,mm 

Fig. 63. PIX. Shock stabilised wall injection. 10" shock. MI = 3.5. 

H = 8.7 MJ/kg. Q = 0.91. 

S 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

- 

H 

- a xx 0 - 

- 8 $ 

- 

- 5d - 

I I I 1111 I 111 Ill I I 

0.1 

0.05 

Fig. 64. PIX. Shock stabilised wall injection. 10' shock. M = 3.5. 

Hs = 6.1 MJ/kg. I 

?- 1 1 I I I I I 

I 

- - 

- - a Q=O 

- + - x, Q = 1.45 

x ++ + - + Q = 1.99 

+++ 

- ?!$:as - 

- + om *Fi - 

OX 

- 

0 

- x I4 

B - 

- 0 

- 

0 

I I I 1111 I Ill Ill I I 

- 

Fig. 65. PIX. Shock stabilised wall injection. 10' shock. 

H = 4.2 MJ/kg , M = 3.5 

S I 

P/Ps 

0.05 

0.025 

.t + 

+ + ++ 

+ 0 

+ ++ 

no 0 0 

0 

I II IIIII IIIIIII I J 

0 7 00 

Fig. 66. PIX. Silane injection from wall. MI = 3.5. 

H = 4.2 MJ/kg S 

U 

+ 

Q=O 

@ = 0.5 

x,mm 
. . . . . . . 000000~ 

I! !! !! !! !! !! ,I 

00n.3 DXt 

_. .

m 

0 

cu 

0 

4 

-?I 

-?I 

0 

0 

0 

a, 

0 

m 

0 

b 

0 

Lo 

0 

m 

0 

v 

0 

m 

T-1 

0 

cu 

0 

-?I 

0 
0000000 

ii II ii ii ii ii 11 

0004 DX+ 

I 

CD ';J m (N t-l 

0 

N 

0 

4 

4 

0 

0 

0 

0 

0 

m 

0 

b 

E 

00 

-A 

0 Ln 

0 

T 

0 

m 

0 

Tu 

0 

4 

0 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

WNV 

000~cuVQD 

. . . . . , , 

OOUQDX+ 

. *o 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I