experimental investigation of mixing and ignition...

EXPERIMENTAL INVESTIGATION OF MIXING AND

IGNITION OF TRANSVERSE JETS IN SUPERSONIC

CROSSFLOWS

a dissertation

submitted to the department of mechanical engineering

and the committee on graduate studies

of stanford university

in partial fulfillment of the requirements

for the degree of

doctor of philosophy

Adela Ben-Yakar

December, 2000

c© Copyright 2001 by Adela Ben-YakarAll Rights Reserved

ii

Abstract

Ignition, flame-holding, and mixing enhancement are fundamental aspects of su-

personic combustion and are critical to the development of hypersonic airbreathing

propulsion engines. High velocities associated with supersonic/hypersonic flight speeds

constrain the performance of propulsion systems because of the limited flow residence

time inside the combustor. A useful hypervelocity propulsion system therefore requires

enhanced mixing of fuel and air, injection with very low drag penalty, and effective

distribution of fuel over the burner cross-section. One of the simplest approaches is

the transverse injection of fuel from wall orifices. The interesting but rather compli-

cated flow-field dynamics of transverse jets injected into a supersonic crossflow has been

studied by many supersonic combustion researchers since 1960’s, but with limited free-

stream flow conditions. Most of the previous research was performed in conventional

wind tunnels by accelerating cold air into supersonic conditions, namely in low velocity

and low total enthalpy flow conditions. However, a real supersonic combustor environ-

ment at flight speeds beyond Mach 8 can only be simulated using impulse facilities due

to the required high total enthalpies. Among various impulse facilities, expansion tubes

are especially useful in providing high total enthalpy flows with the proper chemical

composition, namely the absence of dissociated species.

This research is focused on studying the near-field mixing and ignition properties of

transverse fuel jets injected into realistic supersonic combustor flows. We use advanced

flow visualization techniques, namely planar laser-induced fluorescence (PLIF) imag-

ing of the hydroxyl radical (OH) and ultra-fast-framing-rate schlieren imaging. While

schlieren indicates the location of shock waves, jet penetration and large scale flow

features, OH-PLIF is used to map the regions of ignition.

The first objective of the present work is to characterize the expansion tube facility

iii

for three operating points, simulating flight Mach 8, 10 and 13 total enthalpy conditions.

The ability of the expansion tube to provide a steady-flow test time of adequate duration

and a core flow of sufficient size for 2 mm diameter jet-in-crossflow studies is verified.

The second objective is to study the flow-field properties of hydrogen and ethylene

jets, owing to their relevance to supersonic combustion. Visual observations of image

data, supported by the results for the convection velocity and jet penetration, reveal

significant differences between the hydrogen and ethylene injection cases with similar

momentum flux ratio. Previously the momentum flux ratio was found to be the main

controlling parameter of the jet penetration but the results here demonstrate the exis-

tence of an additional mechanism which alters the vortical structure, the penetration

and the mixing properties of the jet shear layer. The thickness of the penetration band,

used as the representation of the jet-shear-layer thickness is considerable in the ethy-

lene injection case, due to the “tilting-stretching-tearing” mechanism and also due to

the larger growth rate of the jet shear layer. Furthermore autoignition of an ethylene

transverse jet is achieved at flight Mach 10 conditions despite the relatively long ig-

nition delay times of ethylene (hydrocarbons), a key limitation for hydrocarbon-fueled

scramjets. These results of higher penetration, larger jet shear layer growth rate and

autoignition capability indicate that hydrocarbons might be a useful fuel in scramjets

flying at Mach 10 conditions.

The third objective is to investigate the stability of the jet shear layer at various

speed ratios and density ratios via schlieren. The high shear stresses induced by the

large velocity difference across the jet shear layer have a large effect on the structure of

the layer. For the unstable case, we notice: 1) breakdown of Kelvin-Helmholtz structures

with the tilting-stretching-tearing mechanism; 2) increased growth rates with decreasing

values of jet-to-free-stream velocity ratio; 3) large intrusions of crossflow in between the

eddies, and 4) additional shock waves and distortion of the bow shock around the large

eddies. Stable layers show well-defined Kelvin-Helmholtz spanwise rollers. The results

plotted in a density-effective velocity ratio (s − λ) diagram demonstrate two separateregions of “stable”and “unstable”jet shear layers with a separation line at a critical

“effective velocity ratio”.

The final objective is to study the ignition and flame-holding capabilities of a hy-

drogen transverse jet injected into flight Mach 8, 10 and 13 total enthalpy conditions.

The results demonstrate self-ignition in the near-field of the hydrogen jet for the high

iv

total enthalpy conditions (flight Mach 10 and 13). OH-fluorescence is detected along

the jet shear layer periphery in a continuous and very thin filament. For the low total

enthalpy Mach 8 condition, however, the ignition is limited to a small region behind the

bow shock and no OH fluorescence can be observed farther downstream.

It is evident from the results that improved injection schemes for better flame-holding

would be required for practical applications in scramjet engines, especially in the flight

Mach 8 range. During the last few years, cavities have gained the attention of the

scramjet community as a promising flame-holding device, owing to results obtained in

flight tests and to feasibility demonstrations in laboratory scale supersonic combustors.

In this thesis, we summarize the flowfield characteristics of cavities and research efforts

related to cavities employed in low- and high-speed flows. Open questions impacting

the effectiveness of the cavities as flame-holders in supersonic combustors are discussed.

Preliminary studies on cavities with upstream injection are presented indicating self-

ignition inside and around the cavity.

v

Contents

Abstract iii

1 Introduction 1

1.1 Background and Motivation . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.1.1 Typical Scramjet Burner Entry Conditions . . . . . . . . . . . . 1

1.1.2 Flow-Field Features of Jets in Supersonic Crossflows . . . . . . . 4

1.1.3 Ignition and Flame-Holding Strategies in Supersonic Combustion 8

1.2 Thesis Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2 Experimental Aspects 14

2.1 Critical Parameters in Supersonic Combustion Simulation . . . . . . . . 14

2.2 Experimental Facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

2.2.1 Expansion Tube . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

2.2.2 Injection System and its Calibration . . . . . . . . . . . . . . . . 20

2.2.3 Cavity/Injection Plate . . . . . . . . . . . . . . . . . . . . . . . . 23

2.3 Test Flow Characterization in the Flight Mach 8 - 13 Range . . . . . . . 23

2.3.1 Flow Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

2.3.2 Measurement of Flow Properties . . . . . . . . . . . . . . . . . . 27

2.3.3 Boundary Layer Effects on Test Time . . . . . . . . . . . . . . . 32

2.3.4 Core-Flow Size . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

2.3.5 Flow Establishment Time . . . . . . . . . . . . . . . . . . . . . . 37

2.4 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

3 Flow Visualization Techniques 41

3.1 Ultra-Fast Framing Rate Schlieren . . . . . . . . . . . . . . . . . . . . . 41

vi

3.1.1 Previous and Current High Speed Imaging Efforts . . . . . . . . 43

3.1.2 High-Speed Schlieren Imaging Components . . . . . . . . . . . . 45

3.1.3 Timing and Synchronization . . . . . . . . . . . . . . . . . . . . 47

3.1.4 Resolution Considerations . . . . . . . . . . . . . . . . . . . . . . 49

3.1.5 Image Processing and Analysis . . . . . . . . . . . . . . . . . . . 52

3.2 OH-PLIF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

3.2.1 Excitation and Detection Strategy . . . . . . . . . . . . . . . . . 54

3.2.2 OH-PLIF Laser Source and Tuning . . . . . . . . . . . . . . . . . 54

3.2.3 OH-PLIF Imaging System and Its Spatial Resolution . . . . . . . 55

3.2.4 Interpretation of OH-PLIF . . . . . . . . . . . . . . . . . . . . . 56

3.3 Simultaneous Schlieren and OH-PLIF . . . . . . . . . . . . . . . . . . . 56

4 Time Evolution and Mixing Characteristics of Hydrogen and Ethylene

Transverse Jets 59

4.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

4.2 Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

4.2.1 General Flow-Field Features . . . . . . . . . . . . . . . . . . . . 61

4.2.2 Large Scale Coherent Structures . . . . . . . . . . . . . . . . . . 64

4.2.3 Convection Characteristics . . . . . . . . . . . . . . . . . . . . . 74

4.2.4 Penetration and Shear Layer Properties . . . . . . . . . . . . . . 81

4.2.5 OH-PLIF Results . . . . . . . . . . . . . . . . . . . . . . . . . . 84

4.3 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87

5 The Effect of Velocity and Density Ratio on Transverse Jets 88

5.1 Effect of Jet Molecular Weight . . . . . . . . . . . . . . . . . . . . . . . 89

5.1.1 Flow Visualization Results . . . . . . . . . . . . . . . . . . . . . 89

5.1.2 Penetration and Shear Layer Thickness . . . . . . . . . . . . . . 90

5.1.3 Convection Characteristics . . . . . . . . . . . . . . . . . . . . . 95

5.1.4 Characteristic Large Eddy Frequencies (Possible Transverse Jet

Modes) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95

5.1.5 Jet Compressibility Analysis . . . . . . . . . . . . . . . . . . . . 100

5.2 Effect of Density and Velocity Ratios . . . . . . . . . . . . . . . . . . . 104

5.2.1 Flow Visualization Results . . . . . . . . . . . . . . . . . . . . . 104

vii

5.2.2 Definition of an “Effective Velocity Ratio, λ” . . . . . . . . . . . 106

5.2.3 Discussion on the Effect of the Curvature - Centrifugal Instability

Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110

6 Autoignition and Flame-Holding Capability of a Hydrogen Transverse

Jet 114

6.1 Ignition and Flame-Holding Considerations . . . . . . . . . . . . . . . . 114

6.2 Ignition and Flame-Holding Results . . . . . . . . . . . . . . . . . . . . 117

6.2.1 Simultaneous OH-PLIF/Schlieren Results . . . . . . . . . . . . . 117

6.2.2 Top View OH-PLIF Images . . . . . . . . . . . . . . . . . . . . . 117

6.2.3 Comparison of Ignition in Flight Mach 8-13 Total Enthalpy Range 120

6.3 Discussion of the Ignition Process . . . . . . . . . . . . . . . . . . . . . . 123

6.3.1 Ignition Characteristics of Hydrogen . . . . . . . . . . . . . . . . 123

6.3.2 Ignition in Supersonic Combustors . . . . . . . . . . . . . . . . . 126

6.3.3 Ignition of a Hydrogen Transverse Jet . . . . . . . . . . . . . . . 127

6.3.4 Ignition of Ethylene Transverse Jet . . . . . . . . . . . . . . . . . 128

7 Cavity Flame-Holders 132

7.1 Review of Previous Research . . . . . . . . . . . . . . . . . . . . . . . . 132

7.1.1 Cavity Flow-Field Characteristics . . . . . . . . . . . . . . . . . . 132

7.1.2 Cavity in Reacting Flows . . . . . . . . . . . . . . . . . . . . . . 141

7.1.3 Outstanding Questions . . . . . . . . . . . . . . . . . . . . . . . . 148

7.2 Preliminary Cavity Results . . . . . . . . . . . . . . . . . . . . . . . . . 151

7.2.1 Visual Observation of Cavities Using Ultra-Fast Schlieren . . . . 153

7.2.2 Preliminary Ignition Results of Injection/Cavity Schemes . . . . 157

8 Concluding Remarks 161

8.1 Summary of Major Results and Conclusions . . . . . . . . . . . . . . . . 161

8.1.1 Experimental Aspects . . . . . . . . . . . . . . . . . . . . . . . . 161

8.1.2 Flow Visualization Techniques . . . . . . . . . . . . . . . . . . . 163

8.1.3 Characteristics of Hydrogen and Ethylene Transverse Jets . . . . 165

8.1.4 Density and Velocity Ratio Effects . . . . . . . . . . . . . . . . . 166

8.1.5 Ignition and Flame-Holding Capability of a Hydrogen Transverse

Jet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167

viii

8.1.6 Cavity Flame-Holders . . . . . . . . . . . . . . . . . . . . . . . . 167

8.2 Recommendation For Future Work . . . . . . . . . . . . . . . . . . . . . 169

A Expansion Tube Equations 175

B Maps of Estimated Expansion Tube Test Conditions 178

Bibliography 184

ix

List of Tables

1.1 Advantages and disadvantages of expansion tubes relative to shock tun-

nels for hypervelocity combustion simulations. . . . . . . . . . . . . . . . 5

2.1 Test gas (free-stream) flow properties simulating the burner entry condi-

tions of three flight Mach numbers. The corresponding values are from

Fig. 1.1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

2.2 Summary of measured, ideal (inviscid 1-D) and predicted (based on

Mirels solution) properties of test gas for Mach 10 and 13 flow condi-

tions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

4.1 Jet exit flow properties. . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

5.1 The general flow exit properties of gaseous jets with different molecular

weights. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

5.2 The specific flow exit properties of gaseous jets used in the study of the

jet molecular weight effect. The free-stream used in these experiments

simulates the flight Mach 10 flow condition. . . . . . . . . . . . . . . . . 90

5.3 Summary of the different conditions used in the study of jet instability

analysis. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105

7.1 Summary of cavity oscillation frequencies, fm, for different cavity length

to depth ratios, L/D. The table includes the expected values based on

Rossiter’s formula and the ones measured in our experiments. . . . . . . 153

8.1 Recommended free-stream flow conditions for further ignition studies. . 170

x

List of Figures

1.1 Typical scramjet burner entry conditions as a function of flight Mach

number, calculated assuming adiabatic compression. a)The burner en-

try Mach number, M3, for different temperature ratios, T3/T0. b) The

burner entry pressure, p3, and the flight trajectories of constant dynamic

pressure, q0, of 50 and 100 kPa. In our experiments, total enthalpy flows

(Mach number, M3, and static temperature, T3) simulating three nominal

flight conditions (Mach 8, 10 and 13) were generated. . . . . . . . . . . . 2

1.2 Schematic of an underexpanded transverse injection into a supersonic

cross-flow, (a) instantaneous side view at the center-line axis of the jet;

(b) 3-D perspective of the averaged features of the flow-field (Gruber et

al. 1995). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.3 Flow-field schematics of traditional injection/flame-holding schemes for

supersonic combustors. a) underexpanded fuel injection normal to the

crossflow, b) fuel injection at angle, c) injection behind a sudden expan-

sion produced by a step. . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.1 Expansion tube facility (12 m in length and 89 mm inner diameter) and

imaging system. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

2.2 Expansion tube distance-time (x-t) diagram calculated for flight Mach

13 condition. Method of characteristics was used to solve the flow gasdy-

namics properties assuming one-dimensional inviscid theory. Test time

is defined as the time that the test gas has uniform flow quantities and

determined by the time arrival of the contact surface to the tube exit,

and that of the first subsequent rarefaction wave (reflected rarefaction

head in our case of high total enthalpy simulations). . . . . . . . . . . . 19

xi

2.3 Schematic of the test section (27 x 27 cm cross section) where a rake of 4

pitot probes, instrumented with pressure transducers, was located 2.5 cm

downstream of the tube exit. The flow history during the expansion tube

operation was detected via pitot pressure information. Note that the

inner diameter of the tube is 8.9 cm. . . . . . . . . . . . . . . . . . . . . 21

2.4 Optical set-up to measure the test gas velocity, assumed to be equal to

the CS - contact surface velocity. IR emission from 5% CO2 seeded in the

test gas nitrogen is collected by an InSb IR detector at the viewing port

located at 101.6 cm from the end of the tube. The test gas velocity can

then be calculated by considering its time of arrival at the viewing port

and at the pitot rake. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

2.5 Schematic of a) Injection system, b) cavity/injection plate system. . . . . 23

2.6 Schlieren visualization of an underexpanded gaseous injection into still

air. (a)-(c) hydrogen (d)-(e) ethylene jets. The exposure time of the

images was 3musec. Mach disk height, y1, was measured for different

pressure ratios, Pj/Peb, to calibrate the injection system. . . . . . . . . . 24

2.7 Maps of estimated test gas (nitrogen) conditions (state 5) at the exit of an

expansion tube: a) pressure and temperature, b) total enthalpy and ve-

locity of the test gas are plotted for different initial driven and expansion

section pressures. Calculations are performed using the inviscid 1D equa-

tions for a given driver pressure of P4 = 600 psig (helium). Note that the

effective filling pressure of the driver section is taken as P4,eff = 686 psig,

as its inner diameter (10.2 cm) is larger than that of the driven and ex-

pansion sections (8.9 cm). This area difference is accounted for in the

curves presented above. . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

2.8 Example of IR emission, pitot pressure, wall pressure records and the

Mach number variation based on the pitot-to-static pressure ratios, as

a function of time for the Mach 10 flow condition. t = 0 represents

incident shock arrival at the pitot probe, placed 2.5 cm downstream of

the tube exit, while the wall pressure transducer and IR detector are

positioned 40.6 cm and 101.6 cm upstream of the tube exit, respectively

(see Fig. 2.4). Note that the time scale of the static pressure trace is

shifted by 235µs to match the shock arrival at the pitot probe. . . . . . 29

xii

2.9 Example of IR emission, pitot pressure and wall pressure traces as a

function of time for the Mach 13 flow condition. . . . . . . . . . . . . . . 31

2.10 Example of IR emission, pitot pressure and wall pressure traces as a

function of time for the Mach 8 flow condition. . . . . . . . . . . . . . . 32

2.11 Comparison of the measured contact-surface velocity (test gas velocity)

with the shock-induced gas velocity estimated using the measured shock

speeds in the expansion section. . . . . . . . . . . . . . . . . . . . . . . . 33

2.12 X-t diagrams for the expansion section only. 1-D inviscid calculations

are plotted in straight lines and results applying Mirels’ model to include

the boundary layer effects are plotted in dashed lines. One can see the

improved test time as a result of the contact surface (CS) acceleration

due to the developing boundary layer behind the incident shock in the

low pressure expansion section helium flow. The incident shock velocity

was measured and assumed to be constant along the expansion section. 34

2.13 Useful core-flow size in flight Mach 13, 10 and 8 conditions, (a), (b) and

(c), respectively, determined by measuring the radial variation of pitot

pressure at different distances from the tube exit. . . . . . . . . . . . . . 38

3.1 Schlieren imaging set-up. . . . . . . . . . . . . . . . . . . . . . . . . . . 46

3.2 Examples of schlieren images of jet issuing into quiescent air as obtained

for different positions of the knife edge (razor blade) at the focal point.

We use the set-up demonstrated in (d) where the knife edge cuts the

focused light at an angle to enhance both the vertical and the horizontal

density gradient effects. . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

3.3 Timing diagram of the high-speed rate imaging system and its synchro-

nization with the expansion tube test flow time. . . . . . . . . . . . . . . 48

3.4 Examples of schlieren images with different integration/exposure times:

a) 100 ns exposure time, resolving the instantaneous features of the flow-

field, b) 200 ns exposure time, resulting in blurring of the image, c) 3µs

exposure time, averaging the general features while enhancing the weak

shocks such as upstream separation shock wave and downstream recom-

pression wave. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

xiii

3.5 Timing diagram of the high-speed rate imaging system and its synchro-

nization with the expansion tube test flow time. . . . . . . . . . . . . . . 52

3.6 a)Triggering diagram and timing connections of the imaging, the injec-

tion and the data acquisition systems. b) Timing diagram of simultane-

ous OH-PLIF and schlieren and their synchronization with the expansion

tube test flow time. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

4.1 Examples of hydrogen (a) and ethylene (b) injections into a supersonic

crossflow (nitrogen). Exposure time of each image was 200 ns. The x-axis

is normalized by the jet diameter d. . . . . . . . . . . . . . . . . . . . . 62

4.2 An example of schlieren image with 3µs exposure time for hydrogen

injection case. While the unsteady features (coherent structures) are

averaged to zero, some of the weak shocks such as upstream separation

shock wave and downstream recompression wave are emphasized. . . . . 63

4.3 (a)Bow shock position and its angle at the center-line of the jet as mea-

sured from the long exposure schlieren image shown in Fig. 4.2. (b) The

free-stream velocity behind the bow shock and the flow turning angle

based on the measured bow shock shape. For the calculations a calori-

cally perfect gas has been assumed. . . . . . . . . . . . . . . . . . . . . . 65

4.4 An example of 8 consecutive schlieren images of underexpanded hydrogen

injection (d=2mm) into a supersonic crossflow (nitrogen) obtained by

high-speed-framing camera. Exposure time of each image is 100 ns and

interframing time is 1µs. Free-stream conditions are: U∞=2360m/s,

M∞=3.38, T∞=1290 K, p∞=32.4 kPa; and jet-to-free-stream momentum

ratio is: J=1.4±0.1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 694.5 The second example of 4 of 8 consecutive schlieren images of hydrogen

injection into flight Mach 10 condition. Exposure time of each image is

100 ns and interframing time is 1µs. . . . . . . . . . . . . . . . . . . . . 70

4.6 Time evolution of an ethylene jet in a supersonic crossflow (nitrogen)

as observed from 8 consecutive schlieren images. Exposure time of each

image is 100 ns and interframing time is 1.5µs. Free-stream conditions

are: U∞=2360 m/s, M∞=3.38, T∞=1290K, p∞=32.4 kPa; and jet-to-

free-stream momentum ratio is: J=1.4±0.1. . . . . . . . . . . . . . . . . 71

xiv

4.7 The second example of an ethylene transverse jet flow-field in a supersonic

crossflow as observed from 8 time correlated schlieren images. Exposure

time of each image is 200 ns and interframing time is 1.2µs. . . . . . . . 72

4.8 Schematic of the three-dimensional shape (Ω shape) of the unsteady vor-

tical structures formed intermittently (Brizzi et al. 1995). . . . . . . . . 73

4.9 Development of a large-scale ethylene structure (eddy number “-1” in

Fig. 4.7) as it goes through the tilting and stretching processes. Four

different parts of the eddy structure were independently tracked in the

duration of the 8.6µs flow visualization time. . . . . . . . . . . . . . . . 74

4.10 Space-time trajectories of large-scale eddies present in the hydrogen jet

shear layer. The center of the eddies are tracked from the 8 successive

schlieren images shown (a) in Fig. 4.4 and (b) in Fig. 4.5. . . . . . . . . . 75

4.11 Space-time trajectories of ethylene large scale eddies as tracked from 8

time-correlated schlieren images: (a) x-t diagram of the example shown

in Fig. 4.6, (b) x-t diagram of the example shown in Fig. 4.7. . . . . . . 76

4.12 Convection features of coherent large scale structures present in the hy-

drogen jet/free-stream shear layer. The data were subtracted by analyz-

ing the eddy displacement in 8 consecutive schlieren images of 2 exper-

iments (images shown in Figs. 4.4 and 4.5). (a) the convection velocity

of eddies in streamwise and transverse directions, Uc,x and Uc,y, respec-

tively; (b) the convection angle of eddies. . . . . . . . . . . . . . . . . . . 77

4.13 Convection features of eddies present in the ethylene jet/free-stream shear

layer. The data were subtracted by analyzing the eddy displacement in 8

consecutive schlieren images of 2 experiments (images shown in Figs. 4.6

and 4.7). (a) the convection velocity of eddies in streamwise and trans-

verse directions, Uc,x and Uc,y, respectively; (b) the convection angle of

eddies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

4.14 Measured convection velocity of large eddy structures in the hydrogen and

ethylene jet shear layers. The results are compared with the estimated

values of the free-stream velocity immediately behind the bow shock. . . 79

4.15 Schematic showing the low- and high-speed regions of the bow shock-

induced free-stream velocity around the large-scale ethylene eddies. . . . 81

xv

4.16 Transverse penetration data of (a) hydrogen jet and (b) ethylene jet. The

data points were obtained by manually tracking the visually observable

outer edge of the jet from 8 consecutive schlieren images for J = 1.4±0.1.Both of the figures include analysis of 2 experiments namely 16 images.

For comparison, also shown in the figures is the penetration correlation

given by other studies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83

4.17 OH-PLIF results mapping the ignition regions at the jet center-line of:

a) hydrogen injection into air, b) ethylene injection into air, c) ethylene

injection into pure oxygen. . . . . . . . . . . . . . . . . . . . . . . . . . . 86

5.1 Examples of instantaneous schlieren images of jets with different molec-

ular weights. Free-stream conditions are: U∞=2360m/s, M∞=3.38,

T∞=1290K, p∞=32.4 kPa. . . . . . . . . . . . . . . . . . . . . . . . . . 91

5.2 Jet transverse penetration along the axial distance, x/d. Data for four

gases with different molecular weights are presented: a)Mw = 2, J =

1.84, b) Mw = 4, J = 1.72, c) Mw = 8, J = 1.85, d) Mw = 16, J = 1.67.

For comparison, empirical correlations suggested by Gruber et al. (1995)

and Rothstein and Wantuck (1992) are also included for J = 1.75. . . . 93

5.3 Convection velocity of large scale structures in the streamwise (Mc,x)

and transverse (Mc,y) directions as a function of axial distance x/d. The

results for each case (for each molecular weight of jet) are obtained from

4-5 experiments each including 8 consecutive schlieren images. . . . . . . 96

5.4 Formation frequency of the large scale structures and the corresponding

“preferred mode Strouhal number”, Std = fjd/Uj , as a function of the jet

exit velocity. The data were collected from the time evolution observation

of the jet from 8 consecutive schlieren images. Each data point was

obtained by averaging 5-10 experiments with the error bars representing

the deviation from the mean value. . . . . . . . . . . . . . . . . . . . . . 98

5.5 Formation frequency of the large scale structures and the “initial vortex

shedding Strouhal number”, Stθj = fθjθj/Uj , as a function of the jet

Reynolds number. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

5.6 Flow-field schematics used in the jet compressibility analysis. Letters A,

B and C indicate the zones of the jet shear layer. . . . . . . . . . . . . . 101

xvi

5.7 Estimated convective Mach number in zone “A”, MAc , (refer to the

schematic in Fig. 5.6) and the measured visible jet shear layer thickness,

δvis, at x/d≈22 as obtained from penetration width measurements. . . . 1025.8 Estimated velocity fields for the jet and the free-stream in zones “B” and

“C”. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103

5.9 Schlieren images at selected conditions given in Table 5.3. . . . . . . . . 107



5.12 Velocity vector field (U∞, Uj) for a skewed mixing layer and the “effective

velocity ratio”, λ., described in the total velocity vector direction. . . . . 110

5.13 Jet-to-free-stream density ratio vs. velocity ratio. The number near the

data points corresponds to the experimental conditions summarized in

Table 5.3. “Unstable” flow jet is defined when the large structures lose

coherence downstream of the injection port and significant distortions in

the bow shock shape can be observed. . . . . . . . . . . . . . . . . . . . 111

5.14 Jet-to-free-stream density ratio vs. the “effective velocity ratio”, λ. The

number near the data points corresponds to the experimental conditions

summarized in Table 5.3. . . . . . . . . . . . . . . . . . . . . . . . . . . 111

5.15 Schematics illustrating the stability regions based on a) Rayleigh criterion

for centrifugal forces in the curved mixing layers as given in Eq. 5.15 where

cons. = 3 + 2 δvishmax and b) current experimental results. . . . . . . . . . . 112

6.1 Simultaneous OH-PLIF and schlieren images visualizing hydrogen injec-

tion into supersonic crossflow. Free-stream conditions are M = 3.57,

T = 1300K, P = 0.32 atm, V = 2500m/s. The jet-to-freestream mo-

mentum flux ratio is J = 1.4. a) Schlieren image, b) OH-PLIF image

demonstrating the ignition and combustion regions of jet-in-crossflow at

high enthalpy condition, c) Overlaid OH-PLIF and schlieren images. . . 118

6.2 Instantaneous top-view OH-PLIF images obtained at different height

above the injection plate. Free-stream conditions are M=3.57, T=1300K,

P=0.32atm, V=2500m/s. The jet-to-freestream momentum flux ratio is

J=1.4. a) y/d=3, b) y/d=2.5, c) y/d=2, d) y/d=1 above the injection plate.119

xvii

6.3 Instantaneous OH-PLIF acquired at center-line axis of the hydrogen jet

injected into flight Mach 10 and 13 conditions. The images are obtained

by combination of 2 different instantaneous images: near the exit of the

jet (−5 < x/d < 1) and downstream of the jet (1 < x/d < 10). . . . . . . 1216.4 Two instantaneous OH-PLIF images acquired at center-line axis of the

hydrogen jet injected into flight Mach 8 conditions. . . . . . . . . . . . . 122

6.5 Explosion limits of a stoichiometric hydrogen-oxygen mixture (after Sung

et al., 1999). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124

6.6 Variation of ignition delay times τign of a stoichiometric mixture of H2and air with temperature and pressure. Calculations are perfomed using

Chemkin and the GRI mechanism. a) τign vs. T , b) pτign vs. T . . . . . 125

6.7 Variation of ignition time with fuel-air equivalence ratio, φ, for cold H2(Tjet = 300 K) injected into hot air. The values of the ignition delay time

are calculated for different air temperatures, Tair. . . . . . . . . . . . . . 126

6.8 The free-stream temperature and pressure (T2 and P2) behind the bow

shock, measured from schlieren images as discussed in Section 4.2.1 (see

Fig. 4.3). Ignition delay times are calculated for several conditions of air

assuming φ = 0.2. The free-stream flow properties simulate the flight

Mach 10 conditions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127

6.9 Comparison of ignition delay of a stoichiometric mixture of C2H4 (ethy-

lene) and air/oxygen at 1 atm with a stoichiometric mixture of H2 and

air. Two different reaction mechanisms are used to calculate the ignition

delay times of C2H4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129

6.10 a)Variation of ignition delay of a stoichiometric mixture of C2H4 (ethy-

lene) and air/oxygen at various pressures. . . . . . . . . . . . . . . . . . 130

xviii

7.1 Flow-field schematics of cavities with different length to depth ratios,

L/D, in a supersonic flow. a) Open cavity flow for L/D < 7 − 10;shear layer reattaches to the back face while spanning over the cavity.

Small aspect ratio cavities (L/D < 2 − 3) are controlled by transverseoscillation mechanism while in larger aspect ratio cavities longitudinal

oscillation becomes the dominant mechanism. b)Closed cavity flow for

L/D > 10 − 13; shear layer reattaches to the lower wall. The pressureincrease in the back wall vicinity and the pressure decrease in the front

wall results in large drag losses. . . . . . . . . . . . . . . . . . . . . . . . 134

7.2 Typical longitudinal cavity oscillations are caused by the impingement

of the free shear layer on the rear wall which generates travelling shocks

inside the cavity. The shear layer spanning the cavity becomes unsteady

as a result of these acoustic waves deflecting the shear layer up and down,

and/or by the shock induced vortices generated at the front wall leading

edge of the cavity. As a result unsteady waves emanate from the cavity. 135

7.3 Different concepts can be employed to suppress the cavity oscillations:

a)Cavities with an angled back wall suppress the unsteady nature of the

free shear layer by eliminating the generation of the travelling shocks

inside the cavity due to the free-shear-layer impingement. b) In addition,

small disturbances produced by spoilers or by the secondary jet injection

upstream of the cavity can enhance the free-shear-layer growth rate. The

thickening of the cavity shear layer alters its instability characteristics,

such that its preferred roll-up frequency is shifted outside of the natural

frequency of the cavity, and as a result the oscillations are attenuated. . 137

xix

7.4 Instantaneous schlieren images with 200 ns of exposure time demonstrat-

ing the effect of the back wall angle on the flowfield structure of a cav-

ity exposed to a supersonic flow. The free-stream was generated in an

expansion tube to simulate Mach 10 total enthalpy conditions at the su-

personic combustor entry: M∞ = 3.4, U∞ = 2360m/s, T∞ = 1290K,

p∞ = 32 kPa. The boundary layer thickness at the trailing edge of the

cavity is approximately 1mm. a) Cavity with L/D = 5 shows the un-

steady nature of the shear layer at the reattachment with the trailing

edge of the back wall. b)Cavity with slanted back wall (20o) stabilizes

the shear layer reattachment process. . . . . . . . . . . . . . . . . . . . . 138

7.5 Effect of length-to-depth ratio, L/D on a) magnitude (root-mean-square)

of pressure fluctuations on the bottom of the cavity (at x/D = 0.33),

b) drag of the cavity at Mach 1.5 and 2.5 flows. The values were adapted

from Zhang and Edwards (1990). . . . . . . . . . . . . . . . . . . . . . . 139

7.6 Cavity-actuated supersonic mixing enhancement concepts: (a) Sato et al.

(1999), studied the influence of acoustic waves, emitted from a cavity and

impinging on the initial mixing layer. (b)Yu and Schadow (1994) used

the same concept to enhance the mixing of supersonic reacting jets. . . . 143

7.7 Axisymmetric combustor of the Scramjet engine which was flight-tested

by Russian-CIAM/NASA joint program (1998). In this engine two cav-

ities with angled-rear wall were used for flame-holding purposes. The

dimensions are in mm (McClinton et al. 1996). . . . . . . . . . . . . . . 146

7.8 Position of pressure transducers located at the bottom of the cavity to

measure the history of the flow oscillations inside the cavity. Pressure

transducer located farther downstream at x/D = 1.5 provided a more

accurate oscillation frequency measurements. . . . . . . . . . . . . . . . 151

7.9 Examples of cavity pressure traces in arbitrary units: a)L/D = 3, b)L/D =

5, c) L/D = 5 with upstream hydrogen injection, d)L/D = 7. t = 0

represents incident shock arrival at the cavity. The free-stream (N2) con-

ditions represent Mach 10 total enthalpy at the supersonic combustor

entry: M∞ = 3.4, U∞ = 2360 m/s, T∞ = 1290K, p∞ = 32 kPa. . . . . . 152

xx

7.10 Schlieren images demonstrating the differences in the flow-field structure

of cavities with different length-to-depth ratios and back wall angle. The

depth of the cavities is constant and equal to D = 3mm. The free-

stream was generated to simulate Mach 10 total enthalpy conditions at

the supersonic combustor entry: M∞ = 3.4, U∞ = 2360 m/s, T∞ =

1290K, p∞ = 32 kPa. The boundary layer thickness at the trailing edge

of the cavity is approximately 1mm. . . . . . . . . . . . . . . . . . . . . 155

7.11 Schlieren images demonstrating jet interaction with different cavities.

The hydrogen jet is injected into a non-reacting free-stream 3 mm up-

stream of the cavity from a d = 1mm orifice. The injection is performed

at angle of 30o to the plate. The free-stream, N2, represents the flight

Mach 10 burner entry conditions. . . . . . . . . . . . . . . . . . . . . . . 156

7.12 Instantaneous OH-PLIF images demonstrating the ignition regions of a

hydrogen jet interacting with a L/D = 3 cavity. The images are obtained

from 5 single shots at the same conditions. Hydrogen is injected 3mm

upstream the cavity leading edge at an angle of 30o. The free-stream

(air) properties represent the flight Mach 10 burner entry conditions:

M∞ = 3.4, U∞ = 2360 m/s, T∞ = 1290 K, p∞ = 32 kPa. Note that

a schlieren image is also included to indicate the flow-field properties

around the cavity. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 159

7.13 Instantaneous OH-PLIF images demonstrating the ignition regions of a

hydrogen jet interacting with a cavity with L/D = 3 and 30o back wall.

The images are obtained from 5 single shots at the same conditions.

Hydrogen is injected 3 mm upstream the cavity leading edge at an angle

of 30o. The free-stream (air) properties represent the flight Mach 10

burner entry conditions: M∞ = 3.4, U∞ = 2360 m/s, T∞ = 1290 K,

p∞ = 32 kPa. Note that a schlieren image is also included to indicate the

flow-field properties around the cavity. . . . . . . . . . . . . . . . . . . . 160

8.1 Flow-field schematics demonstrating different concepts of angled jet in-

jection combined with cavity flame-holder. a) upstream injection, b) base

injection , c) cavity injection. . . . . . . . . . . . . . . . . . . . . . . . . 171

8.2 Flow-field schematic of a shock-wave/jet interaction. . . . . . . . . . . . 172

xxi

8.3 Schematic of the 25o wedge to generate a shock wave above the injection

plate. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172

8.4 (a)An oblique shock wave impinging the hydrogen jet as visualized us-

ing schlieren imaging. The shock was produced by a 25 o angled wedge

mounted above the injection plate. Flight Mach 13 free-stream condi-

tion. (b)Combined OH-PLIF and schlieren images visualizing the effect

of shock/jet interaction on OH number density. . . . . . . . . . . . . . . 174

B.1 Maps of estimated test gas (nitrogen) conditions at the exit of an expan-

sion tube: a) pressure and temperature, b) total enthalpy and velocity of

the test gas are plotted for different initial driven and expansion section

pressures. Calculations are performed using the inviscid 1D equations for

a given driver pressure of P4 = 300 psig (helium). . . . . . . . . . . . . . 179

B.2 Maps of estimated test gas (Argon) conditions at the exit of an expansion

tube: a) pressure and temperature, b) total enthalpy, velocity and Mach

number of the test gas are plotted for different initial driven and expan-

sion section pressures. Calculations are performed using the inviscid 1D

equations for a given driver pressure of P4 = 300 psig (helium). . . . . . 180

B.3 Maps of estimated test gas (Argon) conditions at the exit of an expansion

tube: a) pressure and temperature, b) total enthalpy, velocity and Mach

number of the test gas are plotted for different initial driven and expan-

sion section pressures. Calculations are performed using the inviscid 1D

equations for a given driver pressure of P4 = 600 psig (helium). . . . . . 181

B.4 Maps of estimated test gas (Helium) conditions at the exit of an expan-

sion tube: a) pressure and temperature, b) total enthalpy, velocity and

Mach number of the test gas are plotted for different initial driven and

expansion section pressures. Calculations are performed using the invis-

cid 1D equations for a given driver pressure of P4 = 300 psig (helium). . 182

B.5 Maps of estimated test gas (Helium) conditions at the exit of an expan-

sion tube: a) pressure and temperature, b) total enthalpy, velocity and

Mach number of the test gas are plotted for different initial driven and

expansion section pressures. Calculations are performed using the invis-

cid 1D equations for a given driver pressure of P4 = 600 psig (helium). . 183

xxii

Chapter 1

Introduction

1.1 Background and Motivation

The success of future hypersonic airbreathing propulsion systems will be largely

dependent on efficient injection, mixing and combustion processes inside the super-

sonic / hypersonic combustion chamber. At flight speeds beyond Mach 6, air entering

the combustor must be supersonic to avoid excessive dissociation of both nitrogen and

oxygen gases. Consequently, the time available for fuel injection, fuel-air mixing and

combustion is very short, of the order of 1 msec, which results in troublesome constraints

on the combustion problem (Ferri 1973; Kumar et al. 1989).

1.1.1 Typical Scramjet Burner Entry Conditions

The combustor entry conditions (Mach number, static temperature and pressure)

of hypersonic airbreathing propulsion systems depend on the flight conditions of the

vehicle. In order to keep the density inside the combustor high for efficient combustion

and the lift at reasonably high values, the flight Mach number, M0, should increase as

the altitude of the vehicle increases.

Residence time is another issue that has to be considered for efficient performance

of a high-speed propulsion system. The air must be compressed in the diffuser in

order to reduce velocities and increase the flow residence time and therefore to allow

a combustor of reasonable length. On the other hand, the reduced velocities at the

combustor entry are restricted by the maximum allowable compression temperature (in

1

CHAPTER 1. INTRODUCTION 2

5 10 15 20 250

1

2

3

4

5

6

7

8

9

10

11

3

2

1

87

T3/T

0=6

Current experiments

1. Mach 8 (3 MJ/kg)

2. Mach 10 (4 MJ/kg)

3. Mach 13 (6 MJ/kg)

M3=1

Burn

er

Entr

yM

ach

Num

ber,

M3

Flight Mach Number, M0

5 10 15 20 25

0

1

2

3

4

5

6

7

8

9

10q

0=50 kPa

q0=100 kPa

p3=1 atm

Burn

er

Entr

yP

ressure

,p

3,atm

Flight Mach Number, M0

20

25

30

35

40

45

Heig

ht,

km

(a) (b)

FIGURE 1.1 Typical scramjet burner entry conditions as a function of flight Mach number, calculated assum-ing adiabatic compression. a) The burner entry Mach number, M3, for different temperatureratios, T3/T0. b) The burner entry pressure, p3, and the flight trajectories of constant dynamicpressure, q0, of 50 and 100 kPa. In our experiments, total enthalpy flows (Mach number, M3,and static temperature, T3) simulating three nominal flight conditions (Mach 8, 10 and 13)were generated.

the range of 1440-1670K (Heiser and Pratt 1994)) to avoid excessive dissociation in

the exhaust flow. These constraints determine the expected values of combustor entry

Mach number and temperature, M3 and T3, respectively.

Considering the above issues, the expected values of flow conditions at the combustor

entrance of an airbreathing propulsion system were estimated and plotted in Fig. 1.1

as a function of flight Mach number, M0. Calculations were performed for different

burner entry temperature to atmosphere temperature ratios (T3/T0) assuming adiabatic

compression (constant total enthalpy) throughout the diffuser according to Eq. 1.1:

T3T0

=1 + γ0−12 M

20

1 + γ3−12 M20

(1.1)

As shown in Fig. 1.1a, for hypersonic flights beyond Mach 6, (M0 > 6) a supersonic

combustion ramjet (scramjet) where the flow remains supersonic / hypersonic through-

out the engine should be considered.

Furthermore, to keep structural loads on the hypersonic vehicle at acceptable levels,

namely, to keep the dynamic pressure, q0 = 1/2ρ0V 20 , in the range of 50-100 kPa, flight

at high speeds is confined to altitudes of 25-40 km. Consequently, the burner entry

pressure, p3 , can be directly evaluated (see Fig. 1.1b) for fixed dynamic pressure, q0 ,


of 50 and 100 kPa, compression efficiency, ηc, of 0.9 and temperature ratio, T3/T0, of 6

using Eq. 1.2 (Heiser and Pratt 1994):

P3P0

=

( T3T0

T3T0

(1− ηc) + ηc

) γcγc−1

(1.2)

where γc is the average specific heat ratio; γc = 1/2(γ0 + γ3). Therefore, at flights

beyond Mach 8, typical pressures at the entrance of supersonic combustors range from

approximately 0.2 to 4 atm depending on the operating parameters for the flight mission,

such as the Mach number and the altitude.

Most supersonic combustion research in the open literature has focused on flight

speeds of Mach 8 and below (Allen et al. 1993; McMillin et al. 1994; Gruber et al.

1995; Parker et al. 1995; Santiago and Dutton 1997), and there are relatively few

works which have been performed for higher flight Mach numbers (Stalker 1989; An-

derson et al. 1990; Bakos et al. 992a; Bakos et al. 992b; Bakos et al. 992c; Erdos

1994; Anderson 1994; Albrechcinski et al. 1995; Wendt and Stalker 1996; Bélanger and

Hornung 1996; McIntyre et al. 1997; Erdos 1998; Ben-Yakar and Hanson 002a). Due

to the large total enthalpies (greater than 3 MJ/kg) associated with high flight Mach

numbers, only impulse facilities are capable of providing the required total tempera-

ture and Mach number to replicate a combustor environment. Expansion tubes and

reflected shock tunnels are two possible types of impulse facilities for ground testing.

Of concern for high stagnation enthalpy simulations is the chemical composition of the

test gas. While in reflected shock tunnels significant amounts of dissociated species are

formed, in expansion tubes the amounts of these species are negligible (?). Therefore,

an expansion tube can provide a more correct simulation of the true flight combustion

chemistry including ignition delay and reaction times. In general, however, expansion

tubes have shorter test times than reflected shock tunnels. The principal advantages

and disadvantages of expansion tubes as compared to other hypersonic test facilities,

especially shock tunnels, are summarized in Table 1.1.

In the present study, the Stanford expansion tube facility is used to generate to-

tal enthalpy conditions in the Mach 8-13 flight range. This facility is one of the few

impulse-type facilities which can provide a wide range of total enthalpies. The free-

piston reflected shock tunnel, T5, located at GALCIT (Bélanger and Hornung 1996),


Calspan reflected shock tunnel (Albrechcinski et al. 1995) and the HYPULSE expan-

sion tube located at GASL (Bakos et al. 992a; Bakos et al. 992b; Erdos 1994) are

three current examples of larger impulse facilities. In these facilities, generic combustor

models with hydrogen injection have been tested using conventional measurement tech-

niques such as pressure measurements along the combustor and flow visualization with

differential interferometry. While most of the high enthalpy and high speed combustion

flow-field studies in the open literature utilize these methods, modern laser-based diag-

nostics can provide flow-field and species information critical for fundamental research

(Erdos 1994; Erdos 1998; Anderson et al. 1992; Rogers et al. 1992; Ben-Yakar and

Hanson 998b).

1.1.2 Flow-Field Features of Jets in Supersonic Crossflows

Efficient performance of very high-speed combustor systems requires fuel and air

mixing at the molecular level in the near-field of the fuel injection. One of the simplest

approaches is the transverse (normal) injection of fuel from a wall orifice. As the fuel

jet, sonic at the exit, interacts with the supersonic crossflow, an interesting but rather

complicated flow-field is generated. Figure 1.2 illustrates the general flow-features of an

under-expanded transverse jet injected into a cross-flow. As the supersonic crossflow

is displaced by the fuel jet a 3-D bow shock is produced due to the blockage produced

by the flow. The bow shock causes the upstream wall boundary layer to separate,

providing a region where the boundary layer and jet fluids mix subsonically upstream

of the jet exit. This region confined by the separation shock wave formed in front of it,

is important in transverse injection flow-fields owing to its flame-holding capability in

combusting situations, as has been shown in previous publications (Parker et al. 1995;

Ben-Yakar and Hanson 998b).

The recent experimental studies performed by Fric and Roshko (1994) provide a

new insight into the vortical structure of a jet injected into a low speed crossflow. Their

photographs, obtained using the smoke-wire visualization technique, illustrate four types

of coherent structures: (1) the near-field jet-shear layer vortices; (2) the far-field counter

rotating vortex pair (CVP); (3) the horseshoe vortex which wraps around the jet column;

and (4) the downstream wake vortices originating from the horseshoe vortex. Figure 1.2

shows the presumed vortical structures for the jet in supersonic crossflow (which are


TABLE 1.1 Advantages and disadvantages of expansion tubes relative to shock tunnels for hypervelocitycombustion simulations.

Shock Tunnels Expansion Tubes

Significant level of radicals such as ”O”and ”NO” are produced in the test gas af-fecting the combustion chemistry. In the re-flected zone of the shock tube, air dissoci-ates due to high temperatures and recom-bines only partially during the fast expansionprocess.

Negligible amounts of radicals are pro-duced. The working gas never stagnates, thusreduces the extent of dissociation. As a resultthe test gas reaches to the test section withmore accurate chemical composition.

The facility needs to contain the to-tal pressures and temperatures of theflow it generates. As noted by Anderson(1994), at flight Mach numbers above 12, thetotal pressure requirements approaches a mil-lion psi or 68,000 atm, which can be producedonly by expansion tubes.

Higher stagnation pressures and tem-peratures can be achieved in expansiontubes even for the same initial driver pres-sure and sound speed, as velocity is addedto the flow through the unsteady expansionprocess without stagnating it.

Free-stream Mach number is fixed bythe nozzle geometry. Simulation of dif-ferent conditions requires replacement of thenozzle with a new geometry.

Variable Mach numbers and conditionscan be easily obtained by just altering theinitial filling pressures.

Nozzle can be damaged due to the highheat transfer rates at the throat and flyingdiagrams inside the tube.

High heat transfer rates are avoided inthe absence of a sonic throat. However, thetest object is prone to damage from flyingdiagrams arriving at the end of the expansiontube operation.

Boundary layer develops throughoutthe nozzle and can be thick compared to thedimensions of the injection port. It is usuallyrequired to eliminate the boundary layer by,for example, inserting a step before the fuelinjection port (Parker et al. 1995)

A thin boundary layer is developed up-stream of the injection port as the injectionplate is placed in the free-jet exiting the tube.

Test times ≈ 1msecTest times are of the order of 1msec. How-ever, a substantial part of it is wasted dur-ing the nozzle start-up time (of the order of0.5msec), required for the supersonic flow tobe established. Test time decreases with in-creasing stagnation enthalpy.

Test times ≈ 0.2 - 0.5 msecNo nozzle start-up time is required. In addi-tion, the establishment of flow on the studiedmodel begins during the expansion sectiongas flow prior to the test gas arrival. As aresult, less useful test time is consumed dur-ing the flow establishment.

Longer test section because of the longertest time and larger core flow. However, sidewall effects should be taken into account.

Test section dimensions depend on thesize of the core flow at the exit of the tubewhich is diminished by the boundary layergrowth on the tube walls.


known to exist in subsonic jet-in-crossflow) as they were partially observed by numerous

studies (Gruber et al. 1996; Gruber et al. 997a; Ben-Yakar et al. 998a).

The origin of the jet vortical structures was studied by several researchers (Fric and

Roshko 1994; Brizzi et al. 1995; Yuan et al. 1999). Among those studies, Yuan et al.

(1999) performed a large-eddy simulation of transverse jets in subsonic crossflows. Their

results revealed that the majority of the jet vortical structures arose from the Kelvin

Helmholtz (K-H) instability of the jet-shear layer in the near-field. Interestingly, they

do not observe the formation of vortex rings around the periphery of the jet as was

assumed in previous studies. Instead they find two kinds of vortices originating from

the jet exit boundary layer: 1) regularly formed spanwise rollers on the upstream and

downstream edges (large scale jet shear layer vortices), 2) quasi-steady vortices, the so-

called ”hanging vortices” that form in the skewed mixing layers (mixing layers formed

from non-parallel streams) on each lateral edge of the jet leading to the formation of

the CVP.

The near-field mixing of transverse jets is dominated by the so-called ”entrainment-

stretching-mixing process” driven by large scale jet-shear layer vortices. In the region

near the injector exit, the injectant fluid moves with a higher velocity tangent to the

interface than the free-stream fluid. As a result, large vortices are periodically formed

engulfing large quantities of free-stream fluid and drawing it into the jet-shear layer

(macromixing). These large scale vortices also stretch the interface between the un-

mixed fluids. Stretching increases the interfacial area and simultaneously steepens the

local concentration gradients along the entire surface while enhancing the diffusive mi-

cromixing.

Preliminary examinations (Gruber et al. 997a; Ben-Yakar and Hanson 002b) of the

convection characteristics of these large-scale structures, developed in a sonic transverse

jet injection into supersonic crossflows, determined that in the far-field the eddies tend

to travel with velocities that are closer to the free-stream velocity. This indicates that

in high speed free-stream conditions, these large coherent structures, where the fuel and

air are mixed by slow molecular diffusion, will also travel at high speeds. Consequently

the combustion process will be mixing controlled.

High mixing efficiency, however, must be achieved in the near-field of the fuel in-

jection for the success of hypersonic propulsion systems. Therefore, it is important to

understand how these structures and their growth rates evolve as flow and jet conditions


(a)

INJECTANT(Hydrogen or

Ethylene)

BOW SHOCK

BARRELSHOCK

RECIRCULATIONZONE

M¥

>1

RECIRCULATIONZONE

BOUNDARY LAYER

SEPARATEDREGION

MACH DISK

LARGE-SCALESTRUCTURES

(b)

3-DBOW SHOCK

M¥

>1

MACH DISK &BARREL SHOCK

AVERAGEJET BOUNDARY

COUNTER-ROTATINGVORTEX PAIR (CVP)

HORSESHOE-VORTEXREGION

FIGURE 1.2 Schematic of an underexpanded transverse injection into a supersonic cross-flow,(a) instantaneous side view at the center-line axis of the jet; (b) 3-D perspective of the av-eraged features of the flow-field (Gruber et al. 1995).


are changed. Two types of fuel are being considered for use in supersonic combustion:

1) hydrogen and 2) hydrocarbon fuels. The large differences in the molecular weights

of these fuels result in a big variation in injection velocities that might lead to a wide

variation in the jet shear layer growth rate and the mixing properties. However, none

of the previous jet penetration studies (Zukoski and Spaid 1964; Schetz and Billig 1966;

Rogers 1971; Rothstein and Wantuck 1992; Papamoschou and Hubbard 1993; Gru-

ber et al. 1995) found any dominant differences between jets with different molecular

weights. Penetration was shown to be dependent primarily on the jet-to-free-stream

momentum flux, J, expressed by:

J =

(ρu2

)jet

(ρu2)∞(1.3)

Most transverse jet-in-crossflow studies were, however, carried out in cold supersonic

flows (namely low velocities) generated in blow-down wind tunnels. The free-stream

temperatures and velocities in these facilities were usually lower than that expected

in a real supersonic combustor environment. Comprehensive studies still need to be

performed to determine the mixing properties of different type of fuels in a relatively

accurate supersonic combustor environment. These observations gave rise for the fol-

lowing question: “is there any other mechanism or controlling parameter which will

alter the large eddy characteristics of the jet shear layer to enhance its near-field mixing

in realistic conditions?”

We were therefore challenged to study the flow features of hydrogen and ethylene

transverse jets exposed to high-speed supersonic free-streams at realistic conditions

leading to high shear levels.

1.1.3 Ignition and Flame-Holding Strategies in Supersonic Combus-

tion

Different injection strategies have been proposed (Billig 1993; Tishkoff et al. 1997;

Abbitt et al. 1993; Hartfield et al. 1994; Riggins et al. 995a; Riggins and Vitt 995b;

Fuller et al. 1998) with particular concern for rapid near-field mixing. These injec-

tion strategies, both flush-mounted injectors and intrusive injectors, typically rely on

the generation of strong streamwise counter-rotating vortices. As a result, mixing is

enhanced both in macro-scale by entrainment of large quantities of air into the fuel


and in micro-scale due to stretching of the fuel-air interface. Stretching increases the

interfacial area and simultaneously steepens the local concentration gradients thereby

enhancing the diffusive micro-mixing. Micro-scale mixing is required for combustion

since chemical reactions occur at the molecular level. However, efficient mixing of fuel

and air does not directly initiate the combustion process.

Ignition and flame-holding in supersonic flows (Huber et al. 1979; Miller 1994; Im

et al. 1998; Sung et al. 1999; Ben-Yakar and Hanson 998b) are two other important

factors that have to be addressed in the design of an injection system. Once the fuel-

air ignition is established, the combustion depends directly on the efficiency of the

mixing. In order for self-ignition (and therefore combustion) to be accomplished in a

flowing combustible mixture, it is necessary that four quantities have suitable values:

static temperature, static pressure, fuel-air mixture, and the residence time at these

conditions. The ignition is considered accomplished when sufficient free radicals are

formed to initiate the reaction system, even though no appreciable heat has yet been

released. When the conditions of spontaneous ignition exist, the distance li at which

it occurs in a medium flowing at a velocity U is: li = Uτi, where τi is the ignition

delay time. As the combustor velocity U becomes larger, the ignition requires longer

distances.

The primary objective of a flame-holder in a supersonic combustion, therefore, is

to reduce the ignition delay time and provide a continuous source of radicals for the

chemical reaction to be established in the shortest distance possible. In general, flame-

holding is achieved by three techniques: 1) organization of a recirculation area where

the fuel and air can be mixed partially at low velocities, 2) interaction of a shock wave

with partially or fully mixed fuel and oxidizer, and 3) formation of coherent structures

containing unmixed fuel and air, wherein a diffusion flame occurs as the gases are

convected downstream.

These three stabilization techniques can be applied in a supersonic combustor in

different ways. One of the simplest approaches is the transverse (normal) injection of

fuel from a wall orifice (see Fig. 1.3a). As the fuel jet interacts with the supersonic

crossflow a bow shock is produced. As a result, the upstream wall boundary layer

separates, providing a region where the boundary layer and jet fluids mix subsonically

upstream of the jet exit. This region is important in transverse injection flowfields owing

to its flame-holding capability in combusting situations, as has been shown in previous


Fuel

Bow Shock

MM¥

>1>1

Fuel

Bigger

Recirculation Region

Weaker Bow Shock

(~ Mach Wave)

Fuel

Smaller

Recirculation Region

Autoignition

Zones

Combined Bow and

Step-Induced Shock

(a)

(b)

(c)

FIGURE 1.3 Flow-field schematics of traditional injection/flame-holding schemes for supersonic combustors.a) underexpanded fuel injection normal to the crossflow, b) fuel injection at angle, c) injectionbehind a sudden expansion produced by a step.

publications (Huber et al. 1979; Ben-Yakar and Hanson 998b; Ben-Yakar and Hanson

999a). However, this injection configuration has stagnation pressure losses due to the

strong 3-D bow-shock formed by the normal jet penetration, particularly at high flight

velocities.

Another way of achieving flame stabilization is by means of a step, followed by

transverse injection (see Fig. 1.3c). The step creates a larger recirculation area with the

hot gases serving as a continuous ignition source. This approach can provide sustained

combustion but, like the previously described method, has the disadvantage of stagna-

tion pressure losses and increase in drag due to the low flow pressure base behind the

step.

On the other hand, it is possible to reduce the pressure losses associated with the


injection process by performing angled injection (e.g., 60o or 30o rather than 90o) so that

the resulting bow shock is weaker (see Fig. 1.3b). In this approach, jet axial momentum

can also contribute to the net engine thrust. Riggins et al. (995a) studied the thrust

potential of a supersonic combustor at Mach 13.5 and Mach 17 flight conditions with

30o flush wall injection of hydrogen and concluded that the major component of thrust

potential gain is due to the jet momentum. In our previous work (Ben-Yakar and

Hanson 998b; Ben-Yakar and Hanson 999a), autoignition of a hydrogen jet transversely

injected into Mach 10-13 flight enthalpy flow conditions was observed in the upstream

recirculation region of the jet and behind the bow shock. However, different experiments

(McMillin et al. 1994) performed for similar geometry but at much lower total-enthalpy

flow conditions showed that ignition occurred only far downstream of the jet. Based

on those observations, angled injection is likely to reduce or eliminate these forms of

autoignition and stabilization especially at flight speeds lower than Mach 10. Therefore,

it is likely that a new technique will be required to obtain autoignition and downstream

combustion stabilization.

In recent years, cavity flame-holders, an integrated fuel injection/flame-holding ap-

proach, have been proposed as a new concept for flame-holding and stabilization in

supersonic combustors (Tishkoff et al. 1997). Cavity flame-holders, designed by CIAM

(Central Institution of Aviation Motors) in Moscow, were used for the first time in a

joint Russian/French dual-mode scramjet flight-test (hydrogen fueled) (Roudakov et al.

1993). Further experiments (Vinagradov et al. 1995; Ortweth et al. 1996; Owens et al.

1998) showed that the use of a cavity after the ramp injector significantly improved

the hydrocarbon combustion efficiency in a supersonic flow. Similar flame stabilization

zones, investigated by Ben-Yakar et al. (998a), have been employed within a solid-fuel

supersonic combustor, demonstrating self-ignition and sustained combustion of PMMA

(Plexiglas) under supersonic flow conditions.

In November 1994, NASA contracted CIAM (Roudakov et al. 1996; McClinton

et al. 1996) to continue exploring the scramjet operating envelope from dual-mode

operation below Mach 6 to the full supersonic combustion mode at Mach 6.5. The

proposed combustor design also included two cavity flame-holders (20 mm in depth by

40mm in axial length and 30mm by 53 mm). The performance predictions obtained by

analytical solutions indicated that these cavities would be quite effective as autoignition

and flame-holding devices. Indeed, the recent flight test of this combustor has been


successfully completed, encouraging further investigation of cavity flame-holders.

It is worth noting that, although there is recent interest in cavity flame-holders for

supersonic combustors, their application in subsonic combustors goes back to the 1950’s.

Probably, the first published investigation of cavity flame-holders is due to Huellmantel

et al. (1957), who studied various shapes of cavities to sustain combustion in low speed

propane-air flames. The main purpose of this thesis is to summarize relevant known

characteristics of cavities in supersonic flows and research efforts related particularly to

cavities employed in low- and high-speed combustors.

1.2 Thesis Objectives

The ultimate objective of this dissertation is to investigate near-field mixing and

flame-holding characteristics of different gaseous fuels such as hydrogen and ethylene

injected normally from a single orifice into a realistic supersonic combustor environ-

ment. We apply advanced non-intrusive flow diagnostic techniques such as Planar

Laser-Induced Fluorescence of OH radicals (OH-PLIF) and schlieren imaging using an

ultra-fast-framing rate digital camera. These techniques and the simulation of high

speed and high temperature free-stream conditions enable unique observations that

were not available in the previous studies. The thesis includes four primary elements:

1. The experimental approach: The goal is to generate a relatively accurate

supersonic burner entry condition, namely a radical-free, high total enthalpy air

flow. An expansion tube is used to generate three nominal free-stream conditions

for flight Mach 8, 10 and 13 regimes. The experimental approach is discussed in

Chapter 2 which includes descriptions of the critical parameters that have to be

considered in the simulation of a supersonic combustor environment, the facility

itself and the measurement techniques. The characterization of the test flow is

then presented summarizing determination of the useful test time, core-flow size

and boundary layer effects, issues that have to be addressed to fully characterize

the flow generated in an expansion tube. The flow visualization techniques are

discussed in detail in Chapter 3.

2. Mixing: In Chapter 4, we study the flow features of hydrogen and ethylene trans-

verse jets exposed to high-speed supersonic free-streams at realistic conditions


leading to high levels of shear. Guided by the observations of these experiments,

we continue in Chapter 5 with a more fundamental study looking into the origin

of the observed phenomena. The outstanding questions that we investigate are:

How do the jet shear layer vortices develop and which parameters control their

stability and coherence? What is the contribution of the jet shear layer vortices

to the near-field mixing? Does the penetration mechanism only depend on jet-to-

crossflow momentum ratio as has been proposed for the last 40 years or is there

any other mechanism leading to higher penetration and better mixing properties?

3. Ignition and flame-holding: The ignition and the flame-holding capabilities of

a hydrogen jet in high total enthalpy flow conditions are presented in Chapter 6.

We study the self-ignition regions in the near-field of the jet in flight Mach 8, 10

and 13 flow conditions using OH-PLIF flow visualization. We also compare the

near-field ignition results of a hydrogen transverse jet with an ethylene transverse

jet at flight Mach 10 conditions.

4. Cavity flame-holders: In Chapter 7, an extensive overview of cavities, which

are considered as a promising flame-holding devices for supersonic combustion, is

presented. Open questions impacting the effectiveness of the cavities as flame-

holders in supersonic combustors are then discussed. Preliminary experimental

results are also summarized. The goal is to study the ignition capability of a jet-

cavity configuration and to observe the differences in the shock wave structures

around cavities as the length-to-depth ratio and the geometry of the cavity back

wall are changed.

Chapter 2

Experimental Aspects

Our experimental approach includes the use of an expansion tube to provide a wide

range of variability in the freestream conditions with relatively accurate chemical compo-

sition. The latter is critical for supersonic combustion studies in the high total enthalpy

flows associated with hypersonic air-breathing propulsion systems.

Efforts are focused on achieving three operating points, simulating flight Mach 8, 10

and 13 total enthalpy conditions at the entrance of a supersonic combustor. The ability

of the expansion tube to provide a steady-flow test time of adequate duration and a

core-flow of sufficient size for 2 mm jet-in-crossflow studies is verified.

In the following sections, the important parameters that must be considered in the

design of a supersonic combustion experiment are discussed and the facility and the test

flow characterization techniques are then summarized. Additional test conditions are

characterized for fundamental fluid mechanical studies and are presented in Chapter 5.

2.1 Critical Parameters in Supersonic Combustion Simu-

lation

An experimental simulation of a supersonic reacting flow requires the replication of

5 parameters (Heiser and Pratt 1994). These simulation parameters including pressure

(p), temperature (T ), velocity (u), characteristic length of the model (L) and gas com-

position (νi) must be manipulated to provide the flight values of certain non-dimensional

parameters such as:

14

CHAPTER 2. EXPERIMENTAL ASPECTS 15

Mach number:

M ∼ u√T

(2.1)

Reynolds number:

Re ∼ ρ√T∼ pL u

T 3/2(2.2)

Damköhler number:

Da ∼ Luτc

(2.3)

Damköhler number represents the ratio of flow residence time, L/u, through the com-

bustor to chemical time, τc, and must be larger than 1 to achieve flame-holding and

a complete combustion process. For flame-holding considerations ignition delay time,

τi, replaces the chemical time in Damköhler number, τc = τi. For a hydrogen-air com-

bustion process, the ignition delay time, varies inversely with pressure because of the

two-body reactions and depends exponentially on temperature. As a result, Damköhler

number can be related to basic parameters in the following form:

Da ∼ pLu · exp(θ/T ) (2.4)

where θ is a characteristic temperature for the ignition time.

Consequently, in order to preserve the values of these three non-dimensional param-

eters it is required to simulate all 5 basic parameters, including temperature, pressure,

velocity, model length and the gas chemical composition. However, it is worth noting

the following point: If the chemical composition of the flow, its velocity and temper-

ature were to be duplicated, then a constant value of the product pL would satisfy

the requirements for simulation of the three non-dimensional parameters. Therefore,

from the standpoint of mixing and flame-holding studies a correct simulation of only 4

parameters is essential: chemical composition, temperature, velocity and the product

pL.

In our experimental approach, we replicate 3 of these 4 parameters: the required

burner entry velocity and burner entry static temperature, u3 and T3, respectively,

according to the values of burner entry Mach number, M3, estimated in Fig. 1.1a. The

use of an expansion tube enables acceleration of the air to total enthalpy conditions


(3-6MJ/kgair) corresponding to the Mach 8-13 flight range, without exposing it to

high stagnation temperatures (3000-6000 K). Therefore, the free-stream contains only

negligible amounts of radicals, produced only by the incident shock wave. The test gas,

first shocked to its maximum temperature (1700-2150 K), is then accelerated and cooled

to the required static temperature (1250-1400 K). Through this unsteady expansion

process, the test gas gains in total temperature and total pressure.

Although in our experiments the free-stream flow composition, Mach number and

static temperature correspond to typical scramjet combustor entrance values, its static

pressure is somewhat below that of actual systems. Table 2.1 summarizes the three

nominal test flow conditions, Mach 8, Mach 10 and Mach 13, achieved in the Stanford

expansion tube facility. Furthermore, since the characteristic length scale in our exper-

iments is small, about 2mm (the diameter of the injection orifice), the parameter pL

is not sufficiently high to replicate a real combustor environment. This might result

in chemical kinetic limitations on the H2 - air ignition and combustion process. On the

other hand, this limitation can be circumvented if an elevated concentration of oxygen

is used in the test gas to increase the collision rates as suggested by Bakos et al. (992b).

Finally, in the current study we have shown that in high-enthalpy flows, ignition of

hydrogen, injected transversely into a free-stream of air, can be achieved in the near-

vicinity of the injector, even at low pL values. Therefore, the ignition will be guaranteed

at higher pressures as the Damköhler and Reynolds numbers increase linearly with pL

in realistic systems.

In conclusion, the most important parameters that have to be replicated for su-

personic combustion studies are chemical composition, temperature and velocity of the

free-stream, and the less important parameter is the product pL. Variation in pressure

affects the ignition time linearly, while variation in temperature has an exponential ef-

fect through the activation energy (and hence characteristic temperature ignition time,

θ) in chemical kinetics.


TABLE 2.1 Test gas (free-stream) flow properties simulating the burner entry conditions of three flightMach numbers. The corresponding values are from Fig. 1.1.

Flight Simulation Mach 8 Mach 10 Mach 13

(1) (2) (3)Initial filling pressuresDriven section, He, psig 300 600 600

(2.17MPa) (4.24 MPa) (4.24MPa)Driven section, 95% N2 +5% CO2, psia 0.45 0.5 0.15

(3.10 kPa) (3.45 kPa) (1.04 kPa)Expansion section, He, torr 70 20 2

(9.13 kPa) (2.67 kPa) (0.27 kPa)Free-stream conditionsTotal enthalpy, MJ/kg 2.9± 0.05 3.9± 0.1 6.2± 0.15Mach number 2.40± 0.03 3.38± 0.04 4.66± 0.07Static temperature, K 1400 1290 1250Static pressure, atm 0.65 0.32 0.04

(65.9 kPa) (32.4 kPa) (4 kPa)Velocity, m/sec (measured) 1800± 20 2360± 25 3200± 50Test time, µsec (measured) 170± 10 270± 10 400± 10Test “slug length”, m (velocity× test time) 0.31 0.64 1.28Establishment length for laminar boundarylayer at L1 = 50 mm, m

0.15 0.15 0.15

Maximum measured recirculation regionlength, L2, (djet = 2mm)

∼ 1.5 djet ∼ 2 djet ∼ 4 djet

Establishment time for the jet upstream re-circulation region, m based on (30−70)×L2

0.09 - 0.21 0.12 - 0.28 0.24 - 0.56

Free-stream Reynolds number at the injec-tion port, Rex = 50mm

29,000 22,000 3,800

Boundary layer thickness upstream of theinjection port, mm

0.65 0.75 1.80

Shock speeds in the expansion section, m/s(measured)

2468 3175 3650

Shock Mach number in the expansion sec-tion

2.44 3.14 3.61

Maximum temperature that the test gas isexposed to, T2, K

1690 1750 2140


Drive

r Sec

tion

Drive

n Sec

tion

Expa

nsion

Sec

tion

Dump Tank

Ampli

fiers

&

Interv

al Co

unter

s

ICCD Camerafor OH-PLIF Imaging

578 x 384 Array

DoubleDiaphragm

8 Cha

nnel

Data

Acqu

isitio

n

Syste

m

KnifeEdge

YM 12

00 N

d:YAG

Lase

r

HD 50

0 Dye

Lase

r

HT 1000

Frequency

Doubler

Long durationXenon Arc

Light Source

DichroicMirror

IMACON 468Ultrafast Framing Camera

for Schlieren Imaging(inc. 8 ICCD modules,

each 576 x 384)

Imag

e Acq

uisitio

n

Comp

uters

Driven/ExpansionDiaphragm

Mirror 2

Mirror 1

FocusingMirror

FIGURE 2.1 Expansion tube facility (12m in length and 89 mm inner diameter) and imaging system.

2.2 Experimental Facility

2.2.1 Expansion Tube

The expansion tube facility with its dedicated lasers and optical arrangement is

schematically illustrated in Fig. 2.1. The tube is 12 m in length (including dump tank)

with an inner diameter of 89 mm, and includes three sections: driver, driven and ex-

pansion. The driver section is filled with high pressure helium gas and is separated

by double diaphragms from the lower pressure driven section, which is filled with the

desired test gas. Mylar film (6.35µm thick) is used as the diaphragm material at the

driven/expansion interface to separate the test gas from low pressure helium gas in the

expansion section.


-1 0 1 2 3 4 5 6 7 80.000

0.001

0.002

0.003

0.004

0.005

0.006

0.007

IR detector

Test

Section

1

1

4 10

10

520

23

4

contact surface

1 Quiescent Test Gas

2 Test Gas Behind Incident Shock

3 Expanded Driver Gas

4 Driver Gas

5 Expanded Test Gas

10 Expansion Gas

20 Expansion Gas

Behind Incident Shock

Expansion

Section

(He)

Driven

Section

(CO2/N

2/O

2)

Driver

Section

(He)

first disturbance arrival

rarefaction tail

rarefaction head

reflected

rarefactio

n head

inciden

t shock

, s1

test time

tim

e,sec

x-distance, m

FIGURE 2.2 Expansion tube distance-time (x-t) diagram calculated for flight Mach 13 condition. Methodof characteristics was used to solve the flow gasdynamics properties assuming one-dimensionalinviscid theory. Test time is defined as the time that the test gas has uniform flow quantitiesand determined by the time arrival of the contact surface to the tube exit, and that of thefirst subsequent rarefaction wave (reflected rarefaction head in our case of high total enthalpysimulations).

The operating sequence of an expansion tube is best represented by the distance-time

(x-t) diagram shown in Fig. 2.2. A run is initiated by bursting the double diaphragms,

which generates a shock wave propagating into the test gas and producing flow of

intermediate velocity with an increased pressure and temperature. The shocked test

gas is then accelerated by an unsteady and constant area expansion process from the

driven section into the lower pressure expansion section, while gaining total temperature


and total pressure. The test gas emerging from the downstream end of the expansion

thus has both a higher stagnation enthalpy and higher effective stagnation pressure than

the shock tube flow from which it originated. Further detail on the operating cycle of

an expansion tube can be found in the review papers of Erdos (1994) and Anderson

(1994).

A square viewing chamber of 27×27 cm cross section is mounted at the exit of theexpansion tube (see Fig. 2.3). A rake of pitot tubes or an instrumented model with the

injection system, is positioned in this test section, which is equipped with an opposed

pair of square (13×13 cm) quartz windows for observation and a fused silica slot on topof the chamber for admission of the vertical laser sheet.

Six piezo-electric pressure transducers are mounted along the driven and expansion

sections for shock speed and wall pressure measurements. An additional transducer,

mounted 20.3 cm downstream of the driven/expansion diaphragms, is used to monitor

the unsteady expansion process at that location.

The expansion section is also equipped with sapphire viewing ports for optical mea-

surements during flow characterization experiments. In those tests, an InSb IR detector

(Judson J-10 InSb equipped by a Perry model 720 amplifier) is mounted at a viewing

port (see Fig. 2.4) to detect the arrival of the test gas (at the viewing port) through

the emission of IR light by small amount of CO2 (5%) seeded into the test gas (nitro-

gen). Also, for flow characterization tests, the injection system is replaced with a pitot

rake consisting of four pressure transducers across the diameter of the tube as shown

in Fig. 2.3. The test gas velocity can then be calculated by considering its arrival time

at the viewing port and at the pitot rake. Data from these sensors are recorded at

1Msample/sec on a PC-based, 8-channel (12-bit) computer-scope. The flow imaging

techniques include Planar Laser-Induced Fluorescence of OH radicals (OH-PLIF) and

schlieren imaging using an ultra-high-speed framing digital camera. Detailed descrip-

tion of these systems and their synchronization with the expansion tube operation are

provided in Chapter 3.

2.2.2 Injection System and its Calibration

The injection system is positioned right at the exit of the expansion tube inside

the test section (Fig. 2.5a). The system consists of a flat plate with an attached high

experimental investigation of mixing and ignition...

Documents