Supersonic Combustion Experimental Investigation at T2 Hypersonic Shock Tunnel

Transcription

1 Supersonic Combustion Experimental Investigation at T2 Hypersonic Shock Tunnel D. Romanelli Pinto, T.V.C. Marcos, R.L.M. Alcaide, A.C. Oliveira, J.B. Chanes Jr., P.G.P. Toro, and M.A.S. Minucci 1 Introduction The aerospace technological products have grown that one cannot conceive of putting payloads (satellites) into Earth orbit or beyond using technologies in operation (rockets carry out solid or liquid fuel). The knowledge required to keep the current launching vehicles is already so high that if the countries do not have a technological support for their own industry, they will depend on of the supplier countries and not have independent capacity sustained. Aerospace vehicle limitations for launching payloads into orbit or beyond require a continuous reduction in size, weight and power consumption of launch vehicles. Some solutions to these challenges require paradigm shifts, new production methods, and new technologies of strategic nature. The requirements of platformslaunched satellites, high performance and reliability, as well as the strict limitations of fuel (reduction of size, weight and power consumption) for launching payloads into orbit or beyond provide the development of hypersonic aircraft using hypersonic airbreathing propulsion based on supersonic combustion. The recent intensification of international efforts to develop hypersonic propulsion system based on supersonic combustion, signals that this is the way of effective access to space in a not too distant future. Therefore, the field of Hypersonic Airbreathing Propulsion basd on Supersonic Combustion, will be essential in the near future for the aerospace industry, and allow the man to build hypersonic planes, to reach other continents in hours and achieve low orbits around Earth. Experimental investigation of the supersonic combustion is being carried out using the T2 Hypersonic Shock Tunnel at the Prof. Henry T. Nagamatsu Laboratory of D. Romanelli Pinto T.V.C. Marcos R.L.M. Alcaide 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 1050 D. Romanelli Pinto et al. Fig. 1 Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics. T2 Hypersonic Shock Tunnel, visible at the left of the T3 Hypersonic Shock Tunnel. Aerothermodynamicsand Hypersonics, Figure 1, which is capable of providing total temperature flows and speed (Mach number) enough to duplicate the environment of high enthalpy and thermochemical characteristics close to those encountered during flight vehicles at high speeds in the Earth atmosphere. The scramjet is the only airbreathing propulsion system to be able to provide the thrust needed efficiently in hypersonic flight [1]. In addition, it has the advantage over rocket engines not lead to oxidizing substance, reducing vehicle weight. To get an idea of structural weight savings this fact, one should bear in mind that the first stage of the Saturn-1, a rocket widely used by NASA, must carry 285ton of liquid oxygen to burn 125ton of RP-1 (a type of highly refined kerosene for rocket). Aerospace vehicles using scramjet have no moving parts, and the scramjet works as follows: slows the flow into the air intake through oblique shock waves until the inlet air reaches the velocity and pressure necessary to ensure that there is combustion. Thus, the gases produced in combustor with its high enthalpy and pressure are expanded in the nozzle. This cycle is then responsible for the thrust generated by the engine. Normally, the fuel chosen for the scramjet is Hydrogen, due to the fast time of ignition and high specific impulse. 2 IEAv T2 Hypersonic Shock Tunnel Ground based test facilities, such as shock tube and hypersonic shock tunnels are the only laboratory devices able to provide total temperature and Mach number sufficient to duplicate the environment of high enthalpy and thermochemical

3 Supersonic Combustion Experimental Investigation 1051 Fig. 2 T2 Hypersonic Shock Tunnel. characteristics close to those encountered during flight vehicles at high speeds in the Earth atmosphere [2]. The T2 Hypersonic shock tunnel [3], Figures 1 and 2, consists of a shock tube (two reservoirs kept at different pressures, separated by a set of two diaphragms, DDS), where a convergent-divergent nozzle is coupled at the end of the low-pressure shock tube. The compressed and heated gas (dry air to simulate flight in the Earths atmosphere) behind the incident shock wave (in the shock tube) is expanded to high speeds and high temperatures (in the divergent nozzle section) to produce hypersonic flow in the test section. The T2 Hypersonic Shock Tunnel used for the present experimental investigation is capable of generating high to low enthalpy hypersonic flow conditions. 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, with adequate throat diameter is used to obtain airflow Mach 7 number in the test section. The different Mach numbers achieved in the

4 1052 D. Romanelli Pinto et al. Fig. 3 Supersonic combustion model installed at the T2 Hypersonic Shock Tunnel test section. 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 Basically, supersonic combustion ramjet (scramjet), Figure 3, is an aeronautical engine that has no moving parts and uses shock waves generated during hypersonic flight, to promote compression and deceleration of atmospheric air. This atmospheric air at supersonic speed, is mixed and burned with a on-board fuel suitable for the production of thrust. Therefore, the combustion process occurs in supersonic regime. When the combustion process occurs in subsonic regime, it is called subsonic combustion or ramjet, the predecessor of the scramjet, which already finds many applications. The total pressure loss that occurs through normal shock wave (which slows the flow in ramjets) makes use of these engines impractical at hypersonic speeds. An important feature of the scramjet is a highly integrated system, where engine and vehicle are indistinguishable. This tight integration is caused by the fact that the front section of the vehicle contributes to the compression of atmospheric air, while the rear contributes to the generation of thrust. The net thrust produced by the scramjet is the difference between the thrust (force that propels the vehicle) generated by the expansion of exhaust gases from the rear of the engine and the total drag (force that resists the movement of the vehicle). These forces may produce thrust to the flight of the vehicle or not depending on the balance of these forces in engine design in question.

5 Supersonic Combustion Experimental Investigation 1053 Thus, aerospace vehicles propelled by scramjets carry only the fuel, usually Hydrogen, using atmospheric air as an oxidant itself by acquiring most of the kinetic energy required to reach Earth orbit during atmospheric flight. As a result of self-propulsive nature of the reactors, they are unable to produce thrust while standing still. The static thrust is zero. Accordingly, they must be accelerated to a speed such that the shock waves produced by the air intake are able to compress the atmospheric air. This velocity is approximately four times the speed of sound, Mach 4, considering scramjet. The conceptual design of the supersonic combustion model, Figure 3, consists by a conical region following by cylindrical section, where, internally, there is the fuel tank. The cylindrical part is tightly integrated with the front and rear in order to reduce drag and weight at hypersonic speeds. At the rear has a cone shape, where the products of combustion are exhausted. The schlieren system was assembled for the evaluation of flow established on the model, as well as for checking the on-board Hydrogen gas fuel injection, during the tests. 4 Supersonic Combustion Results Experimental investigation of supersonic combustion at the T2 Hypersonic Shock Tunnel were performed in medium enthalpy (3000 psi on the driver and 3 atm at driven, respectively), operating in equilibrium interface mode. Schlieren visualization system using high speed camera, Cordin C, was used to observe the injection of fuel (on-board Hydrogen gas) during the test time of Fig. 4 On-board Hydrogen gas injection in supersonic airflow.

6 1054 D. Romanelli Pinto et al. Fig. 5 Time-lapse photography of the combustion produts. the T2 Hpersonic Shock Tunnel, Figure 4. One may observe that the conical attached shock wave at the leading edge of the scramjet model reach the cowl of the inlet. The on-board Hydrogen gas fuel is injected into the atmospheric supersonic airflow. Due to the geometry of the scramjet (axissymetrical) is not possible to visualize the combustion. A time-lapse photography obtained through the integrated camera Nikkon D-1, Figure 5, shows the exhaustion of combustion products during the test of supersonic combustion. No intrusive absorption technique by Diode Laser will be applied in the supersonic combustion experimental investigation. This technique allows to measure the water vapor concentration and temperature of the combustion products. 5 Conclusion Supersonic combustion using an axissymetrical model (conical section following by cylindrical section, where, internally, there is the fuel tank and, at the rear section is a cone shape), has been experimentally investigated at the T2 Hypersonic Shock Tunnel, at the Prof. Henry T. Nagamatsu Laboratory Aerothermodynamics and Hypersonic, at the Institute for Advanced Studies. Schlieren photography, of the injection of on-board Hydrogen gas, and time-lapse photography, of the combustion produts in the exhaust section are presented.

7 Supersonic Combustion Experimental Investigation 1055 References 1. Curran, E.T.: Scramjet Engines: The First Forty Years. Journal of Propulsion and Power 17(6) (November-December 2001) 2. Nagamatsu, H.T.: Shock Tube Technology and Design. In: Ferri, A. (ed.) Fundamental Data Obtained from Shock Tube Experiments, ch. III. Pergamon Press (1961) 3. 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)