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Intake Fuel Injection and Shock Induced Combustion

in a Scramjet Engine Model

Takakage ARAI*, Shigeto MASUDA+ and Fuminori SAKIMA++ Department of Mechanical Systems Engineering, Muroran Institute of Technology

27-1, Mizumoto-cho, Muroran, Hokkaido, 050-8585 Japan arai@mmm.muroran-it.ac.jp

Abstract

This paper presents flows and combustion phenomena in Scramjet engine model. Hypersonic flow was obtained by using a small high enthalpy shock tunnel. At Mach number of about 6.9, Scramjet engine model was set in the test section. At first, the oblique-shock compression process was studied. Gaseous Hydrogen was injected into the supersonic cross flow from intake wall. We conducted Shlieren-image visualization and wall static pressure measurements, self-emission observation. Shlieren-image visualization and self-emission were observed by using CCD camera. We verified the supersonic combustion and shock- induced combustion.

INTRODUCTION

Detonation wave and, more generally, shock-induced combustion ramjets represent an alternative to scramjets as propulsion devices for hypersonic vehicles[1]. A number of formidable tasks must be tackled to achieve satisfactory operation of such scramjets. Especially, it is important that the injected fuel must be more homogeneously mixed with airflow upstream of the intended detonation wave and shock-induced

combustion at a high flight Mach number. As this concept, the length of the combustion chamber is decreased to its minimum. The achievement of this concept is to ensure that combustion chamber is used to combust premixed fuel. This can be achieved by injecting the fuel from the intake allowing mixing to occur upstream of combustor[2] as shown in Fig.1[3]. The task of homogeneously mixing fuel and air in the high-speed, high-enthalpy flow is the most

Fig.1 Concept of Shock induced combustion ramjet or Detonation wave ramjet[3]

* Associate Professor, senior member AIAA Email:arai@mmm.muroran-it.ac.jp + Graduate Student ++ Graduate Student, student member AIAA Copyright Ⓒ 2003 by the American Institute for Aeronautics and Astronautics, Inc. All right reserved

12th AIAA International Space Planes and Hypersonic Systems and Technologies15 - 19 December 2003, Norfolk, Virginia

AIAA 2003-6910

Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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crucial for the developing the supersonic combustion ramjet and the shock induced combustion ramjet. A mixing strategy that has enjoyed considerable success is by organized longitudinal (or axial) vortices, generated either through wall-mounted ramp injector[4] or low-angled wall jet injection[5]. Scramjet testings using the impulse facilities had been undertaken in Australia, University of Queensland and Japan, NAL. In University of

Queensland and NAL Japan, the integrated scramjet models and subscale model were tested to obtain the thrust, pressure measurements, respectively[2,6,7]. And we also have been testing the integrated scramjet model to search the basic problems of scramjet and shock induced combustion in Muroran Institute of Technology[8]. In the present investigation, in order to get the evidence of possibility of shock induced combustion ramjet concept under hypersonic flight

Fig.3 Schematic diagram of experimental apparatus

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conditions, a pre-mixed shock-induced combustion scramjet with a fuel injection at intake was tested under Mach 6.9 condition by using a high enthalpy shock tunnel. The fuel was injected from first ramp (Model B) or second ramp (Model A) in the intake of a 2-D external compression scramjet. The results were compared each other.

EXPERIMENTS Facility The Experiments were conducted in the free piston shock tunnel as shown in Figs.2 and 3 at the Muroran Institute of Technology[9,10]. A conical nozzle for Mach 7 (90 mm of exit diameter) was used. Air was the test gas in cases where combustion was to be permitted and nitrogen in cases where combustion was to be suppressed. Scramjet model tested in the present study was shown in Fig.4, which was an external compression scramjet model with fuel injection from the intake wall. The two dimensional fore-body consists of a 20-deg wedge followed by a 5-deg turn. The fuel was injected from first ramp (Model B) or second ramp (Model A). The location of the cowl could be adjusted both in a stream-wise direction and in a direction

perpendicular to flow. Gaseous Hydrogen or Helium at room temperature were injected through three circular injectors with sonic nozzle (2 mm diameter), and with 65 degrees angle (Model A) and 70 degrees angel (Model B) of injection to the free stream. The equivalence ratio, which was calculated by using the injection pressure and free-stream condition, was about 0.1 to 0.3. Fuel injection during the test time was achieved by means of an injection system by using a piston and fast action valve[7]. The distances from the leading edge of compression surface to the injector was 33 mm (Model B) and 66 mm (Model A),

36

20.0° 25.0°

15.5 11 10 10

50.27

66

170

33

36

170

20.0° 25.0°

50.27

15.5 11 10 10

Model A Model B

Figure 4 Scramjet engine model

Table 1 Performance of free-piston shock tunnel and flow configurations

position free stream

1 2 3

Mach No. 6.9 3.56 3.17 1.39 Static

pressure p kPa

0.42 4.88 7.73 47.9

Static temperature

T K 246 714 817 1617

ρ kg/m3 0.0059 0.0238 0.0330 0.1031

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respectively. The free-stream temperature, pressure and Mach number in the combustion chamber could be approximately 1600 K, 50 kPa and 1.4, respectively by means of oblique shock relations. Test section conditions were established by using measured Pitot pressure, nozzle supply pressure. Flow conditions are presented in Table 1. These values indicated the flight conditions of about 35 km altitude except of the temperature condition. Instrumentation Schlieren photograph measurements were taken by using nano-pulse spark flash with about 30 ns exposure time in the cases of the model with float glass sidewalls. And also, high-speed video camera (max: 19,800 fps) was used to obtain the self-emission and flow properties. From these photographs, dada on shock position, Mack numbers in duct were extracted. Pressure measurements were made by semi-conductor type pressure transducers, which were mounted in the wall of the combustion chamber. The positions of the pressure measurement were at 15.5 mm, 26.5mm, 36.5 mm and 46.5 mm from the leading edge of the cowl.

RESULTS Flow of intake and combustion chamber of scramjet Figure 5 shows the schlieren results of for the upstream section to combustion chamber. The oblique shocks from the leading edges of the intake and the ramp were observed, as should be expected. These two oblique shocks from the intake were focused on and terminated at the edge of the cowl. The reflected cowl shock was also observed in the combustion chamber so that the supersonic flow in the combustion chamber was obtained. Therefore, shock systems were observed in the combustion chamber. Figure 6 show Schlieren photograph of hypersonic flow obtained by the shock tunnel on the Model A case. The flow was felt to right. The test condition simulated about 35 km altitude and Mach 6.9 of flight condition. The equivalent ratio was about 0.1. Two oblique shock waves were observed, and they reflected at the cowl. No emission of light was observed on the case of free stream of nitrogen, on the other hand, the strong radiation of light was observed on the case of free stream of air. Self-emission of ultra-violet wavelength light

Figure 5 Flow visualization result

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Figure 7 shows the self-emission of ultra-violet wavelength light (about 290nm to 330nm). On Model A case, the self-radiation was occurred just after the reflected shock wave at the cowl, while on Model B, the self-radiation was observed after the second oblique shock from the ramp. On Model B, combustion might start in the boundary layer, so it is necessary to protect the external combustion (intake combustion) that the fuel injection plume goes far from wall. Figure 8 show the photographs obtained by a high-speed digital video camera (32,000 fps), which was combined with image intensifier and band pass filter (about 290nm to 330nm). The time interval between frames was about 30 µs. When a shock tunnel started, after second frame, i.e., 60 µs, a luminescence was observed inside a combustion chamber. At the 90 µs (third frame), a strong radiation of light was seen just behind the 3rd oblique shock wave

(a) Nitrogen-hydrogen

(b) Air-hydrogen Figure 6 Self-emission and flow

Injection

Injection

Figure 7 Ultraviolet radical emission

Model A

Model B

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(reflected cowl shock wave). Then, emission was continuously seen inside a combustion

chamber. Pressure distributions in the combustion chamber Figure 9 shows the pressure distributions on the combustion chamber wall. It was found that the pressure in the combustion chamber on Model A

Figure 8 Self-emission obtained by high-speedvideo camera

1

2

3

4

5

6

7

8

9

10

11

Figure 9 Pressure distributions in combustionchamber

(a) Model A

(b) Model B Figure 10 Effect of equivalent ratio on wall pressure distribution

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was larger than Model B. But, the pressure rise caused by supersonic combustion was not observed clearly because of small scale of scramjet engine on the both cases. Figure 10 shows the effect of equivalent ratio on the wall pressure distribution. On φ=0.116 case of Model A, the pressure distribution was different from others. The effect of equivalent ratio on pressure distribution was not so clear because models tested were so small.

CONCLUSIONS To establish the basic concept of shock-induced combustion ramjets and find out some tasks for development of shock-induced combustion ramjets, a fuel injection at intake was tested. The impulse facility of free-piston type shock tunnel was used. The present study is positioned the first step to extract some important problems for development of shock-induced combustion scramjets. It was clear that the intake fuel injection was effective to cause the shock induced combustion in the scramjet engine model, and also fuel injection from the second ramp realized to ignite and burn in the combustion chamber, which we planed.

ACKNOWLEDGMENTS This research was supported in part, by Grant-in-Aid for Scientific Research (C), 13650690, 2002, of Japan Society for the Promotion of Science (JSPS).

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