Pulse Detonation Engine - Seminar Paper

Pulse Detonation Engine

Rocket engines that work much like an automobile engine are being developed at NASA’s Marshall Space Flight Center in Huntsville, Ala. Pulse detonation rocket engines offer a lightweight, low-cost alternative for space transportation. Pulse detonation rocket engine technology is being developed for upper stages that boost satellites to higher orbits. The advanced propulsion technology could also be used for lunar and planetary Landers and excursion vehicles that require throttle control for gentle landings.
 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 igniter—such as a spark plug—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 pump—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 Center’s Advanced Space Transportation Program to dramatically reduce the cost of space transportation.

The main differences between the PDE and the Otto engine is that in the PDE the combustion chamber  is open and no piston is used to com- press the mixture prior to ignition (and also that no shaft work is extracted).
 Instead the compression is an integral part of the detonation, and two of the main advantages of the PDE - the efficiency and simplicity - can be explained by the fact that the combustion occurs in detonative mode. The efficiency of the cycle can be explained by the high level of precompression due to the strong shock wave in the detonation.
Also, the simplicity of the device is a result of the fact that the shock wave - responsible for this compression – is an integrated part of the detonation. Therefore, pre-compression through mechanical devices (e.g., a piston) is not necessary. In this sense the PDE is similar to both the pulse-jet (e.g., the engine used for propulsion of the V-1) and the ram jet engine. But in those two cases the mechanism behind the pre-compression is completely different:
For the pulse-jet the pre-compression is a result of momentum effects of the gases, and is a part of the resonance effects of the engine. The resonance effects are influenced strongly by the external conditions of the engine, and the thrust is drastically reduced at higher speeds (approaching speed of sound). Furthermore, both the specific impulse and the specific thrust are significantly lower  than for turbo-jet or turbo-fan engines. This is due to the fact that the levels of preconditioning that can be obtained through the resonance effects are rather low.
• In the ramjet, pre-compression is obtained through the ram effects as the air is decelerated from supersonic to subsonic. The major drawback with this concept is that the engine is ineffective for speeds lower than around Ma=2.

In the PDE the pre-compression is instead a result of interactions between the combustion and gas dynamic effects, i.e. the combustion is driving the shock wave, and the shock wave (through the increase in temperature across it) is necessary for the fast combustion to occur. In general, detonations are extremely complex phenomena, involving forward propagating as well as transversal shock waves, connected more or less tightly to the combustion complex during the propagation of the entity. 
The biggest obstacles involved in the realization of an air breathing PDE are the initiation of the detonation and the high frequency by which the detonations have to be repeated. Of these two obstacles the initiation of the detonation is believed to be of a more fundamental character, since all physical events involved regarding the initiation are not thorough- ly understood. The detonation can be initiated in two ways; as a direct initiation where the detonation is initiated by a very powerful ignitor more or less immediately or as a Deflagration to Detonation Transition (DDT) where an ordinary flame (i.e. a deflagration) accelerates to a detonation in a much longer time span.

Typically, hundreds of joules are required to obtain a direct initiation of a detonation in a mixture of the most sensitive hydrocarbons and air, which prevents this method to be used in a PDE (if oxygen is used instead of air, these levels are drastically reduced). On the other hand, to ignite an ordinary flame requires reasonable amounts of energy, but the DDT requires lengths on the order of several meters to be completed, making also this method impractical to use in a PDE.
 It is important to point out that there are additional difficulties when liquid fuels are used which generally make them substantially more difficult to detonate. A common method to circumvent these difficulties is to use a pre-detonator - a small tube or a fraction of the main chamber filled with a highly detonable mixture (typically the fuel and oxygen instead of air) - in which the detonation can be easily initiated.
The detonation from the pre-detonator is then supposed to be transmitted to the main chamber and initiate the detonation there. The extra component carried on board (e.g. oxygen) for use in the pre-detonator will lower the specific impulse of the engine, and it is essential to minimize the amount of this extra component.

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).

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.
 For example, it was concluded that a valve controlling the inflow of hydrogen and air is a critical component in the engine. This is also the most significant difference between the engine at FOI and the successful one at AFRL described above. This issue is addressed in the ongoing research at FOI, whose goal is to obtain better understanding of the physical processes involved, and thereby providing efficient design strategies for the PDE.

While real gas effects are important considerations to the prediction of real PDE performance, it is instructive to examine thermodynamic cycle performance using perfect gas assumptions. Such an examination provides three benefits. First, the simplified relations provide an opportunity to understand the fundamental processes inherent in the production of thrust bythe PDE. Second, such an analysis provides the basis for evaluating the potential of the PDE relative to other cycles, most notably the Brayton cycle. Finally, a perfect gas analysis provides the 0framework for developing a thermodynamic cycle analysis for the prediction of realistic PDE performance.
 The present work undertakes such a perfect gas analysis using a standard closed thermodynamic cycle. In the first sections, a thermodynamic cycle description is presented which allows prediction of PDE thrust performance. This cycle description is then modified to include the effects of inlet, combustor and nozzle efficiencies. The efinition of these efficiencies is based on standard component performance.
 Any thermodynamic cycle analysis of the PDE must begin by examining the influence of detonative combustion relative to conventional deflagrative combustion. The classical approach to the detonative combustion analysis is to assume Chapman-Jouget detonation conditions after combustion.
The Chapman-Jouget condition is merely the Rayleigh line analysis limited to sonic velocity as the outlet condition, Shapiro4. Detonation is the supersonic solution of the Chapman-Jouget limited Raleigh analysis, Figure 1. The subsonic Chapman-Jouget solution represents the thermally choked ramjet. To insure consistent handling of the PDE and ramjet, this paper uses Rayleigh analysis for both cycles.
A comparison of the ideal gas Rayleigh process loss was made for deflagration and Chapman-Jouget detonation combustion, Figure 2. The comparison was made for a range of heat additions, represented here by the ratio of the increase in total temperature to the initial static temperature. Four different entrance Mach numbers were also considered. The figure of merit for the comparison is the ratio of the increase in entropy to specific heat at constant pressure. The results show that at the same heat addition and entrance Mach number, detonation is consistently a more efficient combustion process, as evidenced by the lower increase in entropy. This combustion process efficiency is one of the basic thermodynamic advantages of the PDE.

To understand the relative importance of each component efficiency to the ideal cycle analysis, component efficiencies were added one at a time. The first component efficiency added was inlet total pressure recovery. For the inlet component efficiency model, MIL STD 5007D total pressure recoveries were used. To use total pressure recovery as an efficiency index, ideal gas relationships were used to transform the total pressure recovery into its associated process temperatures. These process temperatures were then used to compute a compression efficiency for use in the cycle analysis. The resultant fuel consumption comparison is shown in Figure 7. As both the ramjet and the PDE are experiencing the same component efficiency through the same compression process, no change occurred to the relationship between the cycles. The PDE still exhibits reduced fuel consumption at  all Mach numbers.

The next step in the cycle comparison is to introduce degraded combustor component
efficiencies. In this step, a nominal 90% heat release efficiency was used. The results, Figure 8, are similar to the inlet degraded results in that the PDE still exhibits reduced fuel consumption. As before, both the ramjet and PDE are experiencing similar component losses, so no significant relative change in performance occurs.

For nozzle loss modeling, the generally accepted nozzle gross thrust coefficient, CV, is used. Gross thrust is obtained from the equation:
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 ηe 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 ηe of 90%.
The momentum form of the nozzle efficiency, CV, operates only on the ideal thrust.
The ideal thrust, in turn, is solely dependent on the post-combustion fluid entropy, state 4, and ambient pressure, which together define the ideal, isentropic expansion to state Y, Figure 10.
Therefore, the momentum form, CV, is only dependent on the post-combustion entropy state of the fluid, and independent of the postcombustion fluid energy level, state 4. On the other hand, the energy form of the nozzle efficiency, ηe , operates directly on the energy of state 4, as can be seen by comparing Figure 4 with Figure 10. Expanding from state 4 incurs a significant loss. However, as previously explained, state 4 includes the kinetic energy of the detonation wave which is paid back when the gases expand back to static conditions. Therefore, a state reflective of the actual available energy is required. The state reflective of the actual energy available to the system can be determined by considering the instant the detonation arrives at the end of the chamber. At this instant, the entropy of the system is known to be the same as the CJ entropy because the gases expand to rest isentropically.
The known CJ entropy, combined with the known system enthalpy (h0 + qadd), defines state 4’. In  this way, energy conservation is assured. This analysis assumes a thin detonation wave.
When the energy form of nozzle efficiency is applied to the cycle, the difference between state 4 and 4’ becomes important, Figure 10. Counter intuitively, use of the higher energy state, 4, results in lower performance. This occurs because the expansion from the higher energy state leads to higher entropy generation and lower performance. Use of energy state 4’, rather than state 4, is a more representative cycle point for accurate PDE analysis, as it more appropriately represents the available energy:

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.

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. It is the blow-down process which generates thrust and is directly related to the classical nozzle flow.
Thus, again, it is appropriate to keep the detonation-coupled expansion process (4 - 4’) separate from the flow expansion through the nozzle (4’ - 10’) and to assess nozzle losses against this last expansion.  The simplified analysis given in the previous sections made use of constant nozzle thrust coefficients. Real systems, however, have fixed nozzles, and their performance is a function of nozzle pressure ratio. Nozzle pressure ratio can vary an order of magnitude during the dynamic chamber blow-down process. Detailed integrations of the blowdown process show that energy is conserved.
To correctly enter the dynamic blowdown calculation, the system energy, rather than enthalpy, must be evaluated. A state 4” is defined to be the conserved energy and post-CJ entropy, Figure 10. In the dynamic calculation the energy exiting the PDE can be shown to conserve global enthalpy exactly as state 4’. However, since impulse is a function of velocity and kinetic energy is a function of velocity squared, the dynamically calculated impulse is slightly lower than the effective steady state computation. 
A first principals 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 flowfield were performed  using a new 10 nanosecond duration light source. The complete blowdown 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.

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