25-04-2014, 12:30 PM
pulse detonation engine full report for gentle landings.
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INTRODUCTION
The engine operates on pulses, so controllers could dial in the frequency of the detonation in the "digital" engine to determine thrust. Pulse detonation rocket engines operate by injecting propellants into long cylinders that are open on one end and closed on the other. When gas fills a cylinder, an igniterâ € such as a spark plugâ € is activated. Fuel begins to burn and rapidly transitions to a detonation, or powered shock. The shock wave travels through the cylinder at 10 times the speed of sound, so combustion is completed before the gas has time to expand. The explosive pressure of the detonation pushes the exhaust out the open end of the cylinder, providing thrust to the vehicle.
A major advantage is that pulse detonation rocket engines boost the fuel and oxidizer to extremely high pressure without a turbo pumpâ € an expensive part of conventional rocket engines. In a typical rocket engine, complex turbo pumps must push fuel and oxidizer into the engine chamber at an extremely high pressure of about 2,000 pounds per square inch or the fuel is blown back out.
The pulse mode of pulse detonation rocket engines allows the fuel to be injected at a low pressure of about 200 pounds per square inch. Marshall Engineers and industry partners United Technology Research Corp. of Tullahoma, Tenn. and Adroit Systems Inc. of Seattle have built small-scale pulse detonation rocket engines for ground testing. During about two years of laboratory testing, researchers have demonstrated that hydrogen and oxygen can be injected into a chamber and detonated more than 100 times per second.
NASA and its industry partners have also proven that a pulse detonation rocket engine can provide thrust in the vacuum of space. Technology development now focuses on determining how to ignite the engine in space, proving that sufficient amounts of fuel can flow through the cylinder to provide superior engine performance, and developing computer code and standards to reliably design and predict performance of the new breed of engines.
A developmental, flight-like engine could be ready for demonstration by 2005 and a full-scale, operational engine could be finished about four years later. Manufacturing pulse detonation rocket engines is simple and inexpensive. Engine valves, for instance, would likely be a sophisticated version of automobile fuel injectors. Pulse detonation rocket engine technology is one of many propulsion alternatives being developed by the Marshall Centerâ„¢s Advanced Space Transportation Program to dramatically reduce the cost of space transportation.
EXPERIMENTAL SET-UP
One example of a PDE is shown in Fig 1. This particular engine - which was assembled at one of FOI's (the Swedish Defence Research Agency) departments, Warheads and Propulsion - runs on hydrogen and air and is capable of reaching frequencies up to 40~Hz. The experimental set up is rather simple, basically consisting of a straight tube (in this case with a length of about one metre) in which hydrogen and air is injected, and ignited by an ordinary spark plug. In this experimental engine, the pressure transducers are only used to find out whether the engine operates successfully in detonative mode. This can be seen both by the level of pressure and the speed of propagation of the wave (a detonation in hydrogen air reaches pressures over 20 bar and propagates at around 2,000 m/s). That is, the pressure transducers are used just for the experiments and are not necessary for the operation of the engine. Also shown is a spiral, which, since it helps to induce turbulence in the flow field is known to speed up the transition from flame to detonation. The hydrogen enters the engine through twelve holes of 1 mm. diameter at the edge of a 72 mm. diameter disk at the right end of the engine. The air enters between the central body through which the hydrogen is emerging and the interior walls of the tube.
PRINCIPLE OF THE ENGINE
As the name implies the engine operates in pulsating mode, and each pulse can be broken down to a series of events. The time it takes to complete each of these events puts a limit to the performance of the engine, and the thrust can be shown to be proportional to the frequency and volume of the engine. The events in one cycle are shown schematically in Fig 2, where p0 is the ambient pressure, p1 represents the pressure of the fuel and air mixture, p2 is the peak pressure of the detonation and p3 is the plateau pressure acting on the front plate. As stated above, the thrust of the engine is proportional to the frequency of the engine, and in order to reach acceptable performance levels the indicated cycle has to be repeated at least 50 times per second (depending on the application and the size of the engine).
STATUS
The first experiments on the PDE were done in the beginning of the 1940s, and since then several experiments and numerical calculations have been done. No flying applications have been reported in the open literature, and doubts have been expressed regarding the claimed success of some of the earlier experiments. However, in recent years the PDE has received a renewed interest, and especially in the US work in many different fields related to the PDE has been initiated. One of the most promising efforts is pursued at the Air Force Research Lab (AFRL) at Wright Patterson's Air Force Base headed by Dr. Fred Schauer In that group successful operation of a PDE using hydrogen and air at frequencies at least up to 40 Hz has been demonstrated. In a series of experiments, the proportions between air and hydrogen have been varied from stoichiometric (i.e., where in an ideal combustion process all fuel is burned completely) to lean mixtures. Even at rather lean mixtures the engine is reported to operate in detonative mode and to deliver the expected performance.
This is an indication that the engine could operate on liquid hydrocarbon fuels since those fuels (in a stoichiometric mixture with air) and lean hydrogenair mixtures have similar properties regarding the initiation of the detonation. The PDE at FOI described earlier, did not produce clean detonations propagating over the whole length of the engine. In an effort to improve the situation several parameters were varied: ¢ The length of the mixture chamber. ¢ The shape of the contraction section connecting the air supply to the rest of the engine. ¢ The separation between the contraction section and the beginning of the tube. ¢ The position where hydrogen is introduced. ¢ The position of the spark plug. ¢ In four of the geometries a reed valve was also used, in an attempt to uncouple the engine from the supply systems during the initiation of the detonation. In these cases hydrogen was introduced either upstream or downstream relative to the valve. These changes did not result in a successful, detonative operation of the engine. However, localized peak pressures well above those obtained in detonations, and valuable insight regarding detonations were obtained.
NOZZLE LOSSES
For nozzle loss modeling, the generally accepted nozzle gross thrust coefficient, CV, is used. Gross thrust is obtained from the equation: The ideal gross thrust, Fg,i, is derived from ideal expansion to ambient pressure: Where VY is the ideal velocity of the flow expanded to ambient pressure with no losses. To use nozzle gross thrust coefficient, the energy based thrust equation ( 5 ) must be combined with the basic thrust equation ( 4 ). Substituting the definition of the gross thrust coefficient ( 6 ) results in an expression of the actual exit velocity including losses: In this formulation, the thermodynamic efficiency must not include any nozzle thermodynamic losses as they are included in the CV. Using the above formulation for thrust, the fuel consumption for a real engine was computed, as shown in Figure 9, using a CV of 95%. Once again, the PDE sustains its performance advantage at all Mach numbers. This result differs from the previous work of Heiser and Pratt3. To understand the apparent discrepancy with previous results, an examination of the momentum and energy forms of the nozzle efficiency (CV and e respectively) is necessary. The two nozzle efficiencies are directly related, equation ( 10 ). The state indicated by subscript Y, represents the isentropic expansion from state 4 to ambient pressure. Expansion losses result in the actual nozzle exit velocity, V10, being lower than that possible with isentropic expansion, VY. The lower exit velocity equates to lower kinetic energy and higher temperature in the exhaust stream. The 95% value of CV used for this study equates to an e of 90%.
CRUISE POWER COMPARISONS
To complete this study of PDE performance, reduced power levels meant to represent cruise conditions were evaluated. The cruise fuel-to-air ratios of Figure 3 were used. The resulting fuel consumption is presented in Figure 12. The energy conserved PDE maintained its fuel consumption improvement over the ramjet, although its margins of improvement have diminished. These diminished margins are a direct result of the diminished heat addition. Since the heat addition, i.e. combustion, phase of the cycle provides the PDE its efficiency advantage, its advantage reduces as the heat addition reduces This effect is illustrated in more detail in Figure 13, where the fuel consumption at Mach 3 is examined over a range of heat additions. At the higher heat additions, represented here by the higher levels of specific thrust, the PDE enjoys its highest fuel consumption benefit. As heat addition and specific thrust are reduced, the PDE advantage is reduced until at the low power settings the ramjet enjoys the better fuel efficiency. It should be noted, however, that these power settings are not representative of sustained flight, as vehicle drag will far exceed engine thrust.
DYNAMIC CONSIDERATIONS
In order to make a first order comparison between the PDE and ramjet cycles, the present analysis conserved global enthalpy and tracked the entropy generated by the detonation and by process inefficiencies. However, in order to gain further insight into the detailed operation of the PDE cycle and to apply more suitable component efficiencies, the different phases of PDE operation must be carefully described. Such a thermodynamic description will carefully apply conservation of energy and conservation of enthalpy respectively to the imbedded constant mass and steady flow processes which occur during the PDE cycle. For example, the preceding discussions have used global enthalpy considerations to examine the most appropriate state against which to levy nozzle losses. A similar result can be reached through consideration of the wave dynamics in the chamber: After being processed by the detonation wave, each fluid element is brought to rest relative to the closed end of the detonation tube by an isentropic expansion. The coupled expansion is an inherent part of the detonation which burns the mixture in the chamber in a constant mass process. In contrast, the subsequent blow-down of the detonated gas from the chamber is a quasisteady flow process.
CONCLUSION
A first principal’s comparison was made between the performance of the ramjet and PDE cycles. The PDE was found to out perform the ramjet through Mach 5 for the ideal cycle and for representative component efficiencies. Component efficiencies were applied as inlet recovery, combustion heat release efficiency, and nozzle velocity coefficient. It was shown that conserving global energy provided a more representative basis for the assessment of nozzle loss. Application of the energy form of nozzle performance coefficient to the local CJ state resulted in over prediction of entropy generation during the nozzle expansion process. Finally, the effect of throttle setting was examined. It was shown that at high flight speeds and very low throttle settings, the thermodynamic advantage of the PDE is lost. Shadowgraph visualizations of a pulse detonation engine exhaust flow field were performed using a new 10 nanosecond duration light source. The complete blow down cycle was reconstructed by synchronizing the shadowgraph system with the detonation event. Images of the highest quality were obtained with this new visualization system due to the short pulse of light produced by the source. No smear or distortion of the detonation front and shock waves was observed. The experimental visualizations were then compared to preliminary computational modeling results obtained from a newly developed code at the University of Cincinnati. Very good agreement on the structure and development of the exiting detonation wave was obtained.