Pulse Detonation Rocket Research Articles

A numerical study of the rapid deflagration-to-detonation transition

This paper describes numerically the rapid deflagration-to-detonation transition (DDT) in detail in a high-frequency pulse detonation rocket engine. Different from traditional DDT, reactants are injected into the chamber from near the open end and travel toward the closed end. Previous experiments have implied that the gasdynamic shock by injecting in a confined space and the intensive turbulence generated by the high-speed jet play important roles in the detonation initiation, but explanations of how, when, and where the detonation is generated were not presented clearly due to the limitation of experimental observation. In this work, high-resolution two-dimensional simulations are performed to investigate this process employing a physical model similar to the experimental configuration. A new mechanism manifesting itself as a complicated vortex–flame interaction is found for the flame transition from a laminar to compressible or choking regime. It is discovered that the gasdynamic shock, after reflecting from the end wall, triggers the detonation through the gradient of reactivity with the hot spot formed by the collision of the shock and the flame. A dimensionless criterion defined by the ratio of the acoustic speed to the inverse gradient of the ignition delay time is applied to further describe the spontaneous wave propagation from the perspective of chem-physical dynamics. This criterion quantitatively gives a good prediction of the propagating mode from the subsonic deflagration to a developing detonation, even in such a complex scenario as encountered in this work.

Study on effect of geometry on performance of pulse detonation rocket engines

In order to investigate the effect of the geometry on the performance of pulse detonation rocket engines under valveless self-adaptive working mode, the multi-cycle experiment was carried out on the pulse detonation rocket engines with different mixing sections and obstacles, while the operating frequency was 20 Hz. Gasoline was utilized as fuel, and oxygen-enriched air as oxidizer. Experimental results indicate that the DDT distance and DDT time of PDRE can be shortened by using the mixing section of gradual-expansion structure. The effect of the spiraling grooves and annular grooves on the DDT process is weak. The mixture-based specific impulse of the mixing section with gradual-expansion structure is 4.2%-7.2% lower than that of the mixing section with sudden-expansion structure. The mixture-based specific impulse of PDRE with sudden-expansion structure and Shchelkin spirals is the largest, about 116.4 s.

Study on Effect of Obstacle Shapes on Filling Process in Pulse Detonation Rocket Engine

In order to explore the influence of obstacle shapes on the filling process of pulse detonation rocket engine under valveless self-adaptive working mode, the numerical simulations of two-phase single filling process and multi-cycle experiment were carried out, in which gasoline was utilized as fuel, and oxygen-enriched air as oxidizer. The effects of four kinds of obstacles with different shapes on the filling process of PDRE were studied. The results showed that the existence of obstacles blocks the filling of fuel into the tail of detonation tube, resulting in a great non-uniformity of filling results in the axial and radial directions. Fuel droplets collect near the head and wall of detonation tube, and droplets are seriously blocked by orifice plates. The detonation initiation and stable propagation can be achieved by Shchelkin spirals, annular grooves and spiraling grooves, while the detonation initiation cannot be achieved by orifice plates.

Study on operation and propulsion features of a pulse detonation rocket engine with secondary oxidizer injection

To explore an effective method which can help to reduce the pressure and velocity fluctuations and promote the utilization of kinetic energy in the exhaust products of a pulse detonation chamber, a secondary oxidizer injection method has been proposed and investigated experimentally. A gasoline based valveless pulse detonation system was employed and it operated stably from 10 to 50 Hz. With the extra injection, pressure and velocity fluctuations of the detonation products were reduced dramatically, and a thrust increment up to 29.5% was achieved on the same account of fuel. It provides a new way to promote the propulsion performance of pulse detonation engines beyond the traditional methods, e. g. nozzle and ejector optimization. Moreover, it can be used as a supplement to nozzles, by the combination of a fluidic nozzle, a maximum thrust increase of 75.9% was obtained in this study. The reduction of pressure and velocity fluctuations is believed to provide most of the benefits on thrust improve, and stratified fuel distribution, after-burning, and partial-fill effects are also main reasons especially in higher frequencies.

Experimental research on propulsive performance of the pulse detonation rocket engine with a fluidic nozzle

To explore an effective way for the nozzle design of a pulse detonation rocket engine (PDRE), theoretical and experimental researches have been carried out. The exhaust process and optimal converging/diverging area ratios of a nozzle were theoretically analyzed. Pressure and velocity fluctuations of the exhaust flow were considered to be the major issues. A fluidic solution which employs nitrogen flows in the nozzle throat and divergent section has been tested. The fluidic nozzle is able to adjust the effective area ratios according to the incoming flow. The effective convergent area ratio varied from 2.0 to 2.2, and the divergent one varied from 5.0 to 1.8. It is proved that the fluidic nozzle effectively improved the propulsive performance of the engine. A maximum average thrust increase of 137.8% was obtained in proper conditions.

Study on a simplified double-frequency scheme for pulse detonation rocket engines

To reduce hardware complexity and increase operating frequency, a simplified double-frequency scheme has been developed to obtain high-frequency detonations for pulse detonation rocket engines based on the valved and valveless schemes. This study focuses on characteristics of this simplified scheme. The critical points where detonations transit to deflagrations are considered at different operating conditions to clarify the lower boundaries of stable operations. Experimental results indicate that an ideal buffer zone may not be formed similar to the traditional valved schemes. Comparisons with the valveless scheme without the purge process show that the upper limits of oxidizer percentages are consistent to ensure stable operations, which implies that the present scheme operates potentially as the valveless mode. Detailed studies on the employed solenoid valves and impacts of retonations on the upstream passages demonstrate that solenoid valves are not able to realize intermittent supply control effectively. It is concluded that practical operations utilizing solenoid and rotary valves to control periodic supplies may also operate under a valveless scheme actually with a different upper limit of oxygen percentage in oxidizer.

An improved constant volume cycle model for performance analysis and shape design of PDRE nozzle

An improved constant volume cycle (CVC) model is developed to analyze the nozzle effects on the thrust and specific impulse of pulse detonation rocket engine (PDRE). Theoretically, this model shows that the thrust coefficient/specific impulse of PDRE is a function of the nozzle contraction/expansion ratio and the operating frequency. The relationship between the nozzle contraction ratio and the operation frequency is obtained by introducing the duty ratio, by which the key problem in the theoretical design can be solved. Therefore, the performance of PDRE can be accessed to guide the preliminary shape design of nozzle conveniently and quickly. The higher the operating frequency of PDRE is, the smaller the nozzle contraction ratio should be. Besides, the lower the ambient pressure is, the larger the expansion ratio of the nozzle should be. When the ambient pressure is 1.013 × 105 Pa, the optimal expansion ratio will be less than 2.26. When the ambient pressure is reduced to vacuum, the extremum of the optimal thrust coefficient is 2.236 9, and the extremum of the specific impulse is 321.01 s. The results of the improved model are verified by numerical simulation.

Experimental study on the wall temperature and heat transfer of a two-phase pulse detonation rocket engine

An experimental work was conducted to study the wall temperature and heat transfer on a two-phase pulse detonation rocket engine. Three pressure transducers were employed to validate the fully developed detonation wave in the tube, while six thermocouples were used to measure the increasing temperature of the tube wall and calculate the heat load. The heat flux on the deflagration-to-detonation transition (DDT) section could be in the order of megawatt per square meter, which was 1.27–2.81 times that of the propagation section. The effects of operating conditions, such as the operating frequency and the equivalence ratio, on the heat load were also investigated. A positive correlation was observed between the heat load on the tube wall and the operating frequency, but the increasing trend would slow down gradually with the operating frequency. In addition, as the equivalence ratio increased, the heat load increased initially and decreased afterwards. It is suggested that the fuel-rich conditions should be avoided considering fuel wasting and tube wall over-heating. Moreover, heat transfer enhancement and heat-insulating coating should also be used to sufficiently cool down the engine, especially the DDT section.

Operation of a liquid-fueled and valveless pulse detonation rocket engine at high frequency

To increase the operating frequency and enhance the propulsive performance of a pulse detonation rocket engine (PDRE), a valveless mode with the purge gas based on liquid fuel was developed. Liquid gasoline, oxygen-enriched air, and nitrogen were utilized as fuel, oxidizer, and purge gas respectively. Supply of these propellants was controlled by the periodic pressure oscillations in the detonation tube. The purge gas with the highest supply pressure was filled into the detonation tube first. The fuel and oxidizer were injected after the inner pressure decaying to a relatively low level. This delay and the vaporization of the liquid fuel prevented pre-ignition of the fresh detonable mixture. The PDRE can work stably from 10Hz to 160Hz in this scheme. Fully-developed detonation waves were proved to be obtained successfully at a maximum frequency of 130Hz. Both oxygen volume fraction and supply pressure of the purge gas were tested to determine the limits for stable operation. The propulsive performance of a practical PDRE at a wide operating frequency range was also tested and the maximum thrust was obtained at 140Hz. Compared to the results without purge gas, introduction of a purge gas extends the ranges of supply pressure and operating frequency.

Impact of nozzles on a valveless pulse detonation rocket engine without the purge process

To investigate the impact of exit nozzles on the operation and the propulsive performance of a valveless pulse detonation rocket engine (PDRE) without the purge process, experiments have been carried out under wide operating frequencies. Gasoline and oxygen-enriched air were used as fuel and oxidizer. Two detonation tubes and twenty-one nozzles with different shapes were employed in this study. It was observed that both the contraction ratio and the expansion ratio have an influence on the performance of the PDRE. Nozzles with converging sections decrease the highest operating frequency of the valveless PDRE, even if it cannot operate normally with nozzles whose contraction ratios are larger than four. The expansion ratio almost has no impact on the operation of the valveless PDRE, but the thrust decreases when the expansion ratios of diverging nozzles are larger than five. Most converging nozzles increase the thrust of the valveless PDRE, especially when the fill fractions were greater than one, while most diverging nozzles increase the thrust when the fill fractions were smaller than one. A maximum thrust increase of 25% has been obtained when nozzles were utilized.

Analysis of influence design parameters ejector pipe at the detonation specifications engine

The engines that use a detonation cycle have greater thermal efficiency than engines with constant-pressure combustion cycle. The mechanisms of ejector augmentation are evident in secondary air entrainment. The impulse characteristics of a pulse detonation engine (PDE) with ejector were investigated under different parameters. The propagation of a detonation wave in an ejector-augmented pulse detonation rocket engine fueled with a hydrogen-oxygen methane-oxygen and hydrocarbon-oxygen mixture is studied. The straight cylindrical ejector was coaxially installed at different axial locations relative to the exit of the detonation chamber. The use of a non-steady ejector is modeled for pulse detonation engine performance improvement. In this paper, a two-dimensional numerical model is developed to help design the basic geometry and operating parameters of the device. The unsteady flow processes are simulated and compared with a baseline PDE without ejector enhancement. Various features including detonation-shock interaction, detonation diffraction and vortex formation are observed. It is observed that the design of detonation wave flow path in detonation tube, ejector and operating parameters such as Mach numbers are mainly responsible for improving the propulsion performance of PDE.Object оf numerical simulation is the cylindrical detonation chamber with a diameter of Dc = 25 mm and a length of Lc = 100 mm. The cylindrical ejectors with a length of Le = 100 mm and diameters De = 35, 40, 45, 50 that 55 mm were attached to the chamber. Also ejectors with a diameter of De = 35 mm and lengths of Le = 50, 100, 150, 200 that a 250 mm respectively were attached to the chamber. The calculation was performed with a computer program that implements a numerical TVD-scheme. The results of calculations were presented as the dependence of specific thrust impulse Isp/Ispo ratio (Isp ­- specific thrust impulse with ejector nozzle; Ispo - specific thrust impulse of a chamber only) to the ejector-to-PDE diameter ratio (De/Dc), to the ejector-to-PDE length ratio (Le/Lc) and to the displacement of an ejector-to-PDE length ratio ((Ld +Lc)/Lc), where Ld – distance from the input section of the nozzle to the thrust wall of a detonation chamber. The method of a numerical simulation proved the possibility of a significant increase of pulse detonation engine performance through the use of ejector nozzle.

Flight Validation of a Rotary-Valved Four-Cylinder Pulse Detonation Rocket

A rotary-valved four-cylinder pulse detonation rocket engine system, Todoroki II, was developed, in which two novel techniques, the use of an inflow-driven motor and an inverted oxidizer cylinder, were introduced. The total length of the system was 1910 mm; its total weight when filled with ethylene–nitrous-oxide propellant and helium purge gas was 32.5 kg; and the engine weight was 9.6 kg. In a ground firing test with a duration of 1500 ms, a thrust-to-engine-weight ratio of 2.7 was achieved. Thus, it was demonstrated that a multicylinder pulse detonation rocket engine system can be used as a practical thrust mechanism. Using a launch and recovery system, a flight-simulating test was conducted to evaluate the features and viability of the engine design. The launch and recovery system operated perfectly, and Todoroki II reached a height of about 9.7 m. The operation of the pulse detonation rocket engine under conditions simulating real vertical flight without constraint forces with a duration of about 1200 ms and a thrust-to-engine-weight ratio of 2.5 was demonstrated. No serious impact of the vibration caused by the pulse detonation rocket engine operation or the rotation of the rotary valve on the flight was observed.

Operation of a pulse detonation engine system at high frequency

In order to reduce the run-up distance for deflagration-to-detonation transition process, and increase the operating frequency of a pulse detonation rocket engine, a specific experimental system was designed. A tube closed at one end and open at the other end was employed as the detonation tube. The detonable mixtures were injected into the detonation tube periodically in the middle of the tube through a rotary valve and ignited near the open end. Detonation waves can be obtained when the deflagration waves propagate to the closed end. Varieties of experiments were carried out in the pulse detonation rocket engine system utilizing ethylene as fuel and oxygen-enriched air as oxidizer. Fully developed detonation waves were obtained, and a steady operating frequency of 200 Hz was successfully achieved. Higher operating frequency can be expected when the supplying flow rate is improved or smaller tube is utilized.

Thrust Performance of Rotary-Valved Four-Cylinder Pulse Detonation Rocket Engine

Thrust Performance of Rotary-Valved Four-Cylinder Pulse Detonation Rocket Engine

Propulsive performance of a pulse detonation rocket engine without the purge process

To measure the propulsive performance of a high-frequency PDRE (pulse detonation rocket engine), an experimental facility was established. Utilizing the valveless mode, the PDRE was operated without the purge process at a maximum operating frequency of 110 Hz successfully. In this study, oxygen-enriched air instead of oxygen was utilized as oxidizer and liquid gasoline was used as fuel because its vaporization would cool the hot combustion products, which would create a buffer zone and ensure stable operation without the purge process. The thrust under a wide range of operating frequencies was obtained including over and partially filled cases. Based on the results of partially filled conditions, one empirical formula for the partial fill effect was developed. In addition, analysis was carried out which resulted in a quite similar equation. Therefore, a partial fill model for the valveless PDRE without the purge process was obtained. Finally, comparisons between the proposed model and other models developed in traditional operations were performed.

Experimental Study of Ejectors’ Effect on the Performance of Pulse Detonation Rocket Engine

Experimental studies were performed to investigate the ejector’s effect on the performance of a pulse detonation rocket engine (PDRE). The PDRE employed in the experiments utilizes gasoline as fuel and air as oxidizer. The operating frequency of the PDRE is 10 and 20 Hz respectively. Performance is quantified by thrust measurements, and the average thrust of the PDRE is obtained by integration of the thrust measured in the experiment. The experimental results show that the PDRE can increase thrust augmentation with an ejector, a maximum thrust augmentation is measured by 18.7% at 10 Hz operating frequency. In the experiment, axial placement (X/DPDRE) of the ejector is varied downstream the PDRE tube of 0.5,1 and 2 respectively. It is found that the thrust augmentation obviously increases when the relative position at 1 and 2, and also the thrust augmentation increases at higher operating frequency. The experiment result is useful for the development of a PDRE producing higher thrust.

Study on a liquid-fueled and valveless pulse detonation rocket engine without the purge process

In traditional PDRE (pulse detonation rocket engines), mechanical valves are used for periodic supply control and a purge process is widely used to form a buffer zone to prevent fresh fuel-oxidizer mixture from pre-ignition. However, both of them increase hardware complexity and limit increase of operating frequency. To eliminate mechanical valves and the purge process, a valveless mode without the purge process has been proposed. Multi-cycle detonations are able to create periodic pressure oscillations inside a detonation tube, which are capable to interrupt fuel and oxidizer supply. In the present study, liquid gasoline was used because vaporization of the liquid fuel would cool hot combustion products, which acted as a buffer zone. Therefore, a liquid-fueled and valveless PDRE without the purge process became possible. When oxygen-enriched air with 25% ∼ 45% oxygen by volume was employed, a maximum operating frequency of 110 Hz was achieved. It was observed that supply pressures of fuel and oxidizer were of great importance for such an operating mode. Exhaust plumes at different operating frequencies were also investigated. The results indicated that it was feasible for the valveless PDRE to run steadily without the purge process when liquid gasoline was utilized.

Experimental investigations on effects of wall-temperature on performance of a pulse detonation rocket engine

To study the effects of wall-temperature on performance of a pulse detonation rocket engine (PDRE), long-duration experiments were conducted. An internally grooved semi-circle spiral, instead of traditional Shchelkin spiral, was utilized to facilitate the deflagration to detonation transition (DDT) process and prolong run time in the present study. Four pressure transducers and a type K thermocouple were employed to collect data during the experiments. Experiments were carried out at three different operating frequencies, such as 10Hz, 20Hz and 30Hz, until unstable operation happened. The results showed that wall-temperature increased faster at higher operating frequency. It was observed that DDT time and DDT distance both decreased with the increase of wall-temperature, and detonation pressure increased sharply with wall-temperature first and then decreased slowly when wall-temperature exceeded a certain value. Both of them were believed to be due to the impact of hot-wall on fuel vaporization and reactant temperature. Unreasonably high wall-temperature was found to cause pre-ignition of fuel–oxidizer mixture, and this was why unstable operation of the PDRE happened after long-duration runs. In addition, effects of wall-temperature on specific impulse was also discussed.

Thrust Estimates of a Partially Filled Multi-Cycle Pulse Detonation Rocket Engine

Partially filled multi-cycle operation usually happens in pulse detonation rocket engines (PDREs). Thrust estimates of PDREs under such operating conditions will provide significant reference for relevant studies. An analytical model for thrust estimates is proposed in the present study. This model is based on Wintenberger' model for impulse calculation and empirical formula for specific impulse prediction of partially filled tubes; moreover, it takes into account performance penalty created by obstacles preliminarily, which are introduced to accelerate deflagration to detonation transition. Four specific analytical models are developed according to three previously proposed empirical formulas and a fitting one proposed for partial filling effect. Comparisons between model predictions and experimental measurements are carried out to validate the reliability of the models. It is found that the models make reasonable estimates and one of the models performs very well over a wide range of conditions. Discussion and analysis on the proposed model and partial filling effect are also performed in this study. Although there are many other factors that should be considered in evaluating thrust of partially filled multi-cycle PDREs, the present study supplies a rapid and effective means for thrust estimation and provides some modeling ideas meanwhile.

Experimental studies on rotary valves for single-tube pulse detonation rocket engines

Two types of rotary valve were designed and fabricated for single-tube pulse detonation rocket engines. According to their driving components, they were named as gear-driven and cam-driven rotary valve, respectively. Preliminary experimental investigations were carried out to test their feasibility for supply of the single-tube pulse detonation rocket engine. Based on the gear-driven rotary valve, the single-tube pulse detonation rocket engine operated at 10 Hz stably. When operating frequency was increased to 15 Hz, inconsecutive detonations happened which was due to signal output block in the encoder. It was found that optimal ignition timing needed to be adjusted to obtain stable operation in the gear-driven rotary-valved pulse detonation rocket engine. Discussion on this was conducted. Operations of the cam-driven rotary-valved pulse detonation rocket engine were also performed. Fully developed and successive detonations at 40 Hz were successfully obtained although an explosion happened in initial tests. Experimental results showed that these two types of rotary valve were able to realize supply control of single-tube pulse detonation rocket engines effectively.