Missiles that can fly at hypersonic speeds for many minutes have a future Naval need to increase standoff distances and reduce the time to impact on time critical targets. To be feasible these missiles must use air breathing propulsion systems that depend upon both subsonic and supersonic combustion to produce thrust for part of their flight. It is further envisioned that future missile systems will be required to fly at speeds that cover a wide range of Mach numbers. For example a single vehicle's speed might vary from Mach 3-5, Mach 4-7, or Mach 4-8 during a single flight. Mach is the ratio of a vehicle's speed to the speed of sound in the fluid at the local conditions of pressure and temperature. The speed of sound in air, in the atmosphere at sea level, is generally about 1,225 kilometers per hour (1117 ft/sec). A vehicle flying at this speed is said to be flying at Mach 1. Twice this speed is Mach 2 and so on.
Supersonic flight is deemed to be anything between Mach 1 and about Mach 4, or four times the speed of sound. Hypersonic speeds lie above that. The fastest currently existing manned air-breathing jet, known as the SR-71 Blackbird, flies at about Mach 3.6. Hyper-X also known as the X-43 recently flew at almost 10 times the speed of sound, or Mach 10. Speeds over Mach 5 are commonly termed hypersonic. (The Aviation History On-line Museum & GE Aircraft Engines). To achieve such a wide range of flight Mach numbers future vehicles will be required to use both subsonic and supersonic combustion to produce thrust. Turbine based engines can be used to fly from 0 up to approximately Mach 4. A ramjet that has been boosted to around Mach 2.3 can be used to fly up to approximately Mach 5. To attain speeds above Mach 5 future vehicle will probably have to rely on scramjet propulsion. A ramjet engine construct operates by subsonic combustion of fuel in a stream of air compressed by the forward speed of the aircraft itself, as opposed to a normal jet engine, in which the compressor section (the fan blades) compresses the air. Ramjets operate from about Mach 2.3 to Mach 5.
U.S. patent application Ser. No. 10/337,667 filed on Dec. 24, 2002, illustrates the following on scramjet engines. Scramjet is an acronym for Supersonic Combustion Ramjet. The scramjet differs from the ramjet in that combustion takes place at supersonic air velocities through the engine. It is mechanically simple having a burner (2), but vastly more complex aerodynamically than a jet engine. Hydrogen is the ideal fuel used, however other fuels such as hydrocarbons can be used. A ramjet has no moving parts and achieves compression of intake air by the forward speed of the air vehicle. Air entering the intake of a supersonic aircraft is slowed to subsonic velocities and compressed by aerodynamic diffusion created by the inlet and diffuser (1) to velocities and pressures comparable to those in a turbojet augmentor. After fuel injection and combustion the hot gases are accelerated through a nozzle to generate push (thrust).
A scramjet engine construct (supersonic-combustion ramjet) is a ramjet engine in which combustion of fuel in the engine takes place at supersonic velocities. The scramjet has an inlet (1), burner (2), and nozzle (3). Scramjet technology is challenging because only limited testing can be performed in ground facilities. A scramjet works by taking in air at speeds greater than Mach 5, slowing the air velocity to lower supersonic speeds and using it to combust a fuel, accelerating the products of combustion in a supersonic nozzle which in turn creates thrust. Hypersonic missiles will have to utilize both ramjet and scramjet technologies and constructs during a single flight to reach both high speeds and long-range capabilities. In addition transition between the different modes of combustion will be required. Ground test facilities will be required that can simulate variable Mach number air flows to test these future vehicles. Ground test facilities typically depend upon combustion heated air accelerated through a high-speed free jet nozzle to simulate flight like conditions. The ability to vary the free jet Mach number by 1 or 2 Mach numbers during a single test does not currently exist for these high temperature (>1200K) flows. The ability to vary the Mach number during a test is the application for the device described herein. Due to a wide range of flight conditions encountered by these engines during operation, the air mass flow varies considerably while the missile is changing speed and altitude. Changing Mach number and angle of attack necessitate changes in fuel burn rate to maintain the variable fuel consumption within acceptable limits.
Combustion instability has been a problem of major concern. Unstable, periodic fluctuation of combustion chamber pressure that has been encountered in ramburners arises from several causes associated with combustion mechanism, aerodynamic conditions, real or apparent shifts in fuel-to-air ratio or heat release, and acoustic resonance. The periodic shedding of vortices produced in highly sheared gas flows has been recognized as a source of substantial acoustic energy for many years. For example, experimental studies have demonstrated that vortex shedding from gas flow restrictors disposed in large, segmented, solid propellant rocket motors couples with the combustion chamber acoustics to generate substantial acoustic pressures. The maximum acoustic energies are produced when the vortex shedding frequency matches one of the acoustic resonances of the combustor. It has been demonstrated that locating the restrictors near a velocity antinode generated the maximum acoustic pressures in a solid propellant rocket motor, with a highly sheared flow occurring at the grain transition boundary in boost/sustain type tactical solid propellant rocket motors.
An apparatus and method for controlling pressure oscillations caused by vortex shedding is disclosed in U.S. Pat. No. 4,760,695 issued to Brown, et al. on Aug. 2, 1988. The '695 patent discloses an apparatus and method for controlling pressure oscillations caused by vortex shedding. Vortex shedding can lead to excessive thrust oscillations and motor vibrations, having a detrimental effect on performance. This is achieved by restricting the grain transition boundary or combustor inlet at the sudden expansion geometry, such that the gas flow separates upstream and produces a vena contracta downstream of the restriction, which combine to preclude the formation of acoustic pressure instabilities in the flowing gas stream. Such an inlet restriction also inhibits the feedback of acoustic pressure to the point of upstream gas flow separation, thereby preventing the formation of organized oscillations. The '695 patent does not present a method or apparatus, which attempts to permit a significant portion of the required enthalpy proportioned to the expansion side of the nozzle via supersonic combustion without the use of expensive film cooled nozzles. Furthermore, the '695 patent does not utilize an oxygen injection means for maintaining flame stability.
Creating long-duration hypersonic flight simulation conditions in a ground test facility can lead to material problems in the facility hardware. The conventional approach to creating these high Mach, high enthalpy flows for testing engine propulsion systems is to expand very high temperature combustion products through a converging-diverging nozzle to the desired pressure, temperature, and Mach number. However, the high total temperature and pressure required for testing may exceed the material capabilities especially at the throat of the nozzle in the ground test facility. As a result, the conventional high temperature subsonic combustion and nozzle expansion approach requires the use of complex and expensive film cooled nozzles (estimated to cost between $1-2 million) for 100 lbm/sec flow rates to survive the high enthalpy flow conditions for the relatively long test times required by the use of such a device. U.S. patent application Ser. No. 10/337,667 filed on Dec. 24, 2002, illustrates the following and is hereby incorporated by reference. This device used supersonic combustion to generate the high Mach number flows required for testing while overcoming the problems associated with survival of the facility nozzle throat.
Therefore, using the device described in U.S. patent application Ser. No. 10/337,667, the present invention fulfills the need to develop a supersonic combustion heater that can vary the Mach numbers at the exit plane to enable mode transition in future air breathing propulsion systems to be tested. The device described herein accomplishes this variable Mach number by heat addition through a plurality of strategically positioned fuel injections means, enhances kinetics, produces an increased high enthalpy flow source, enhances flame stability, improves mixing between fuel and air, and shortens chemical ignition delay, without the use of expensive film cooled nozzles.
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