Supersonic Combustion Flow Visualization at Hypersonic Flow

Transcription

1 Supersonic Combustion Flow Visualization at Hypersonic Flow T.V.C. Marcos, D. Romanelli Pinto, G.S. Moura, A.C. Oliveira, J.B. Chanes Jr., P.G.P. Toro, and M.A.S. Minucci 1 Introduction Currently, a new generation of scientific aerospace vehicles, using advanced hypersonic airbreathing propulsion based on supersonic combustion technology, is in development at several research centers [1]. The 14-X Brazilian Hypersonic Aerospace Vehicle, Figure 1, designed by Rolim et al. [2], at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Figure 2, at the Institute for Advanced Studies (IEAv), is part of the continuing effort of the Department of Aerospace Science and Technology (DCTA), to develop a technological demonstrator using: i) waverider technology to provide lift to the aerospace vehicle, and ii) scramjet technology to provide hypersonic airbreathing propulsion system based on supersonic combustion. Fig X Hypersonic Aerospace Vehicle. T.V.C. Marcos D. Romanelli Pinto G.S. Moura A.C. Oliveira J.B. Chanes Jr. P.G.P. Toro M.A.S. Minucci Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics Institute for Advanced Studies Rodovia dos Tamoios km 5, São José dos Campos, SP(BR)

2 1042 T.V.C. Marcos et al. Fig. 2 T2 Hypersonic Shock Tunnel, visible at the left of the T3 Hypersonic Shock Tunnel. Aerospace vehicle using waverider technology obtains lift using the shock wave, formed during supersonic/hypersonic flight throught the Earths atmosphere, which originates at the edge and it is attached to the bottom surface of the vehicle, generating a region of high pressure, resulting in high lift and low drag [3]. Atmospheric air, pre-compressed by the shock wave, which lies between the shock wave and the leading edge of the vehicle may be used in hypersonic propulsion system based on scramjet technology. Hypersonic airbreathing propulsion, that uses supersonic combustion ramjet (scramjet) technology [4], offers substantial advantages to improve performance of aerospace vehicle that flies at hypersonic speeds through the Earth s atmosphere, by reducing on-board fuel. Basically, scramjet is a fully integrated airbreathing aeronautical engine that uses the oblique/conical shock waves generated during the hypersonic flight, to promote compression and deceleration of freestream atmospheric air at the inlet of the scramjet. Fuel, at least sonic speed, may be injected into the supersonic airflow just downstream of the inlet. Right after, both oxygen from the atmosphere and on-board fuel are mixing. The combination of the high energies of the fuel and of the oncoming supersonic airflow the combustion at supersonic speed starts. Finally, the divergent exhaust nozzle at the afterbody vehicle accelerates the exhaust gases, creating thrust. 2 IEAv T2 Hypersonic Shock Tunnel The development of such airbreathing propulsion system requires experiments, those may be done in Hypersonic Shock Tunnel (pulsed hypersonic wind tunnel) [4], which reproduces the flight conditions encountered in a scramjet engine in flight. The IEAv T2 Hypersonic Shock Tunnel [5], Figure 2, used for the present investigation, is capable of generating high to low enthalpy hypersonic flow conditions.

3 Supersonic Combustion Flow Visualization at Hypersonic Flow 1043 In the high and medium enthalpy runs, helium is used as the driver gas and the tunnel operates in the equilibrium interface condition to produce a useful test time of roughly 500 s to 1.5 ms, reservoir conditions of 5,000 K to 1685 K and 120 bar to 173 bar, respectively. In the low enthalpy case, air is used as the driver gas to produce a useful test time of about 1.5 ms and reservoir conditions of 950 K and 25 bar. The test section airflow Mach number is 6.2 and 7.3 in the high and medium enthalpy tests, respectively, and 7.8 in the low enthalpy ones. In the present investigation dry air was used as a test gas. Conical nozzle, 15 degree half angle and adequate throat to obtain airflow Mach number 7 in the test section. The different Mach numbers achieved in the test section are the result of the different reservoir conditions and real gas effects present in the tests. The test conditions did not vary more than 5% from run to run. 3 Supersonic Combustion Model The conceptual design of the supersonic combustion model, Figure 3, was based on the configuration of existing compression ramp at the inlet of scramjet of the 14-X Hypersonic Aerospace Vehicle, Figure 1. The 15 degree compression ramp was made of stainless steel. A piezoelectric Pressure Transducer, as a pitot pressure, was installed right above of the ramp. Therefore, the detached normal shock established ahead of the pitot pressure will have no influence in the oblique shock wave established at the edge of the supersonic combustion model. Fig. 3 Supersonic combustion model installed at the T2 Hypersonic Shock Tunnel test section. 4 Supersonic Combustion Flow Visualization Results The supersonic combustion model (2D wedge) allows flow visualization of the combustion phenomenon through the non intrusive schlieren technique. The fuel tank volume was adequate to produce equivalence factor of 1.6. The conventional schlieren arrangement was adopted and conceived to illuminate most of the test-section optical windows. The test section has two 8-in. diameter

4 1044 T.V.C. Marcos et al. quartz windows with a 190-mm. effective vision area. Because of the positions of the optical tables in the laboratory, an auxiliary optical mounting was necessary to redirect the collimated lamp beam to/from the test section and to make it possible to fix the source light at the parabolic-mirror focus. The schlieren visualization system is composed of a xenon flash lamp, two parabolic mirrors (8-in. diameter and 64-in. focal length), a knife edge (razor blade), an ultrahigh speed camera (Cordin 550) used to record the events, and an objective lens. The camera has 32 CCDs and can capture at a maximum rate of 2,000,000 frames per second. In the present work, the camera was set to operate at 100,000 frames per second. For tightness and to qualify the device used to inject on-board Hydrogen gas fuel into Mach number 7 airflow, Helium gas was inject in the quiescent gas at 80 mbar (test section in vacuum), Figure 4, and no hypersonic airflow. The sequence of schlieren photographs of the Mach number 7 airflow over supersonic combustion model, Figure 5, shows the oblique shock wave attached to the leading edge of the model as expected. Observe that there is no gas injection in the Mach number 7 airflow. Also, oblique shock angle of 22 degree agrees with the theoretical calculation by oblique shock theory for freestream Mach number 7 airflow. Fig. 4 Helium gas injection in vacuum and no hypersonic airflow. Fig. 5 Sequence of schlieren photographs of the Mach number 7 airflow over supersonic combustion model and no on-board Hydrogen gas injection.

5 Supersonic Combustion Flow Visualization at Hypersonic Flow 1045 Fig. 6 Schlieren photography of on-board Hydrogen gas injected into Mach number 7 airflow over supersonic combustion model. The schlieren photography of the on-board Hydrogen gas injection into Mach number 7 airflow shows a small increase in the slope of the shock wave after the region of Hydrogen gas injection, Figure 6, which may be the result of heat release, resulting in a decrease of Mach number in the region. However, since the schlieren photographs obtained during the test with on-board Hydrogen gas injection were quite similar to the schlieren photographs of the Mach number 7 airflow over supersonic combustion model without gas injection, an obstacle was added in the supersonic combustion model to create a high temperature stagnant region to increase the probability to start the combustion. Again, the sequence of schlieren photographs of the Mach number 7 airflow over supersonic combustion model, Figure 7, shows not only the 22 degree oblique shock angle attached to the leading edge of the model, but also the interaction of the compressed supersonic airflow, by the oblique shock wave, and the obstacle (located downstream of the Hydrogen gas injection). Fig. 7 Sequence of schlieren photographs of the Mach number 7 airflow over supersonic combustion model, with no gas injection.

6 1046 T.V.C. Marcos et al. Fig. 8 Sequence of schlieren photographs of the Mach number 7 airflow over supersonic combustion model and the injection of the on-board Hydrogen gas. As expected, after place the barrier (obstacle) in front of the supersonic airflow, resulting in a stagnation point and in a significant rise in temperature. The exact moment when the on-board Hydrogen gas injected into airflow the schlieren photographs, Figure 8, show the rapid expansion of mixed gas. Also, this sequence of schlieren photographs shows the time when the shock wave is formed and it is destroyed by the combustion, of the mixed Hydrogen gas and the Oxygen from the supersonic airflow. 5 Conclusion Supersonic combustion using a 2D model (wedge), based on the compression ramp of the 14-X Brazilian Hypersonic Aerospace Vehicle, has been experimentally investigated at the T2 Hypersonic Shock Tunnel, at the Prof. Henry T. Nagamatsu Laboratory Aerothermodynamics and Hypersonics, at the Institute for Advanced Studies. Schlieren photographs, of the combustion of the mixed Oxygen from the supersonic airflow and the on-board Hydrogen gas, show the destruction of the oblique shock wave right at the time when the on-board Hydrogen gas injected meets Oxygen from the supersonic airflow. References 1. Curran, E.T.: Scramjet Engines: The First Forty Years. Journal of Propulsion and Power 17(6) (November-December 2001) 2. Rolim, T.C., Minucci, M.A.S., Toro, P.G.P., Soviero, P.A.O.: Experimental Results of a Mach 10 Conical-Flow Derived Waverider. In: 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, AIAA (2009)

7 Supersonic Combustion Flow Visualization at Hypersonic Flow Wang, Y., Zhang, D., Deng, X.: Design of Waverider Configuration with High Lift-Drag Ratio. Journal of Aircraft 44(1), (2007) 4. Heiser, W.H., Pratt, D.T., Daley, D.H., Mehta, U.B.: Hypersonic Airbreathing Propulsion. AIAA Education Series, 594 p. AIAA (1994) 5. Nagamatsu, H.T.: Shock Tube Technology and Design. In: Ferri, A. (ed.) Fundamental Data Obtained from Shock Tube Experiments, ch. III. Pergamon Press (1961) 6. Nascimento, M.A.C.: Gaseous Piston Effect in Shock Tube/Tunnel When Operating in the Equilibrium Interface Condition. Doctoral Thesis. Instituto Tecnológico de Aeronáutica - ITA, São José dos Campos, SP, Brazil (October 1997) (in English)