Performance enhancement of a pulse detonation rocket engine

Utilizing liquid kerosene as the fuel, oxygen as oxidizer and nitrogen as purge gas, a series of multi-cycle detonation experiments was conducted to improve the performance of pulse detonation rocket engine (PDRE). In order to improve the performance of the engine, it is crucial to develop an effective DDT enhancement device with less flow loss and higher survival in hostile detonation tube; therefore, three spiraling internal grooves were tested. The three spiraling internal grooves were semicircle, square and inversed-triangle grooves, respectively. The results showed that the spiraling internal groove can effectively enhance DDT and prolong the operation time of PDRE. The effect of groove shape on thrust enhancement of PDRE and the optimum length of spiraling groove were then investigated. To improve the detonability of liquid kerosene and prolong the durability of PDRE, experiments on the kerosene preheating based on active cooling were conducted. The results demonstrated that with the aid of fuel preheating, the detonation initiation time for liquid kerosene was noticeably reduced and a fully-developed detonation wave was achieved in the position away from igniter 4.67 times the diameter of the detonation tube. By adding the additive to liquid kerosene, the detonation initiation time from 0.75 ms decreased to 0.34 ms and the detonability of fuel was dramatically improved. Finally, experiments were conducted to investigate the effects of the operating frequency on the detonation parameters, the fill fraction and PDRE performance. The results indicated that detonation pressure and temperature vary with the operating frequency of PDRE, and the fill fraction has a significant influence on the specific impulse of PDRE. With the strategy of partial filling in detonation tube, the specific impulse can be remarkably enhanced.     

Mixing and detonation structure in a rotating detonation engine with an axial air inlet

High-fidelity simulations of an experimental rotating detonation engine with an axial air inlet were conducted. The system operated with hydrogen as fuel at globally stoichiometric conditions. Instantaneous data showed that the detonation front is highly corrugated, and is considerably weaker than an ideal Chapman–Jouguet wave. Regions of deflagration are present ahead of the wave, caused by mixing with product gases from the previous cycle, as well as the injector recovery process. It is found that as the post-detonation high pressure flow expands, the injectors recover unsteadily, leading to a transient mixing process ahead of the next cycle. The resulting flow structure not only promotes mixing between product and reactant gases, but also increases likelihood of autoignition. These results show that the detonation process is very sensitive to injector design and the transient behavior during the detonation cycle. Phase-averaged statistics and conditionally averaged data are used to understand the overall reaction structure. Comparisons with available experimental data on this configuration show remarkable good agreement of the predicted reacting flow structure.     

Critical condition of inner cylinder radius for sustaining rotating detonation waves in rotating detonation engine thruster

We describe the critical condition necessary for the inner cylinder radius of a rotating detonation engine (RDE) used for in-space rocket propulsion to sustain adequate thruster performance. Using gaseous C₂H₄ and O₂ as the propellant, we measured thrust and impulse of the RDE experimentally, varying in the inner cylinder radius r_i from 31 mm (typical annular configuration) to 0 (no-inner-cylinder configuration), while keeping the outer cylinder radius (r_o = 39 mm) and propellant injector position (r_inj = 35 mm) constant. In the experiments, we also performed high-speed imaging of self-luminescence in the combustion chamber and engine plume. In the case of relatively large inner cylinder radii (r_i = 23 and 31 mm), rotating detonation waves in the combustion chamber attached to the inner cylinder surface, whereas for relatively small inner cylinder radii (r_i = 0, 9, and 15 mm), rotating detonation waves were observed to detach from the inner cylinder surface. In these small inner radii cases, strong chemical luminescence was observed in the plume, probably due to the existence of soot. On the other hand, for cases where r_i = 15, 23, and 31 mm, the specific impulses were greater than 80% of the ideal value at correct expansion. Meanwhile, for cases r_i = 0 and 9 mm, the specific impulses were below 80% of the ideal expansion value. This was considered to be due to the imperfect detonation combustion (deflagration combustion) observed in small inner cylinder radius cases. Our results suggest that in our experimental conditions, r_i = 15 mm was close to the critical condition for sustaining rotating detonation in a suitable state for efficient thrust generation. This condition in the inner cylinder radius corresponds to a condition in the reduced unburned layer height of 4.5–6.5.     

The dynamics of a non-premixed rotating detonation engine from time-resolved temperature measurements

The structure and dynamics of a hydrogen-air rotating detonation engine (RDE) are described based on 100-kHz laser absorption spectroscopy measurements of water temperature at four simultaneous locations within the detonation channel. The analysis focuses on the evolution of the flowfield over a 200 ms period for three separate air mass flow rate cases. Two-dimensional unwrapped visualizations of the temperatures show a flowfield structure containing regions with the detonation front, combustion products, oblique shock, and refilling reactants, qualitatively agreeing with previous simulations and experiments. A major conclusion is that water from the combustion products is measured throughout all space and time in the RDE, including near the injector, implying the presence of performance loss processes such as burning upstream of the detonation wave or the back recirculation of combustion products with fresh fuel–air. By analyzing the elevated temperatures of the reactants during the refill process, one estimation for the mass fraction of combustion products in the reactants is as high as 20–30% on average. This product mass fraction is found to be inversely proportional to the bulk air mass flow rate and decreases as time progresses. This indicates these non-ideal processes are more significant closer to RDE ignition for poorer performing operating conditions. For the largest air mass flow case, water temperatures near the nominally cold plenum conditions likely corroborate the presence of a recirculation region on the RDE inner body. Analysis of inter- and intra-cycle temperature dynamics further support non-ideal processes occurring behind the detonation wave and during the refill process. As a whole, the data indicates that the RDE performance is better as time progresses away from ignition or for higher air mass flow rates. These data are also important for comparison with numerical models.     

Numerical simulation of the operation process and thrust performance of an air-breathing pulse detonation engine in supersonic flight conditions

Multidimensional simulations of the unsteady gasdynamic flow in the duct of an air-breathing pulse detonation engine (ABPDE) operating on propane gas and the flow around it in supersonic flight at Mach numbers M of 3.0 and an altitude of 9.3 and 16 km are performed. It is shown that, at a length and diameter of the duct of 2.12 m and 83 mm, respectively, an ABPDE with an air intake and a nozzle can operate in a cyclic mode at a repetition frequency of 48 Hz, with a rapid deflagration-to-detonation transition (DDT) occurring at a distance of 5–6 combustion chamber diameters. To determine the thrust performance of the ABPDE in flight conditions, a series of working cycles were simulated with consideration given to the external flow around the engine. Calculations showed that the specific impulse of the ABPDE is approximately 1700 s. This value is much higher than the specific impulse typical of ramjet engines operating on conventional combustion (1200–1500 s) and substantially lower than the specific impulse obtained for the atmospheric conditions at sea level at zero flight velocity (∼2500 s).     

Numerical simulation of a mixed-mode reaction front in a PPC engine

The ignition process, mode of combustion and reaction front propagation in a partially premixed combustion (PPC) engine running with a primary reference fuel (87% iso-octane, 13% n-heptane by volume) is studied numerically in a large eddy simulation. Different combustion modes, ignition front propagation, premixed flame and non-premixed flame, are observed simultaneously. Displacement speed of CO iso-surface propagation describes the transition of premixed auto-ignition to non-premixed flame. High temporal resolution optical data of CH₂O and chemiluminescence are compared with simulated results. A high speed ignition front is seen to expand through fuel-rich mixture and stabilize around stoichiometry in a non-premixed flame while lean premixed combustion occurs in the spray wake at a much slower pace. A good qualitative agreement of the distribution of chemiluminescence and CH₂O formation and destruction shows that the simulation approach sufficiently captures the driving physics of mixed-mode combustion in PPC engines. The study shows that the transition from auto-ignition to flame occurs over a period of several crank angles and the reaction front propagation can be captured using the described model.     

Three-dimensional numerical thrust performance analysis of hydrogen fuel mixture rotating detonation engine with aerospike nozzle

The propulsive performance for an H₂/O₂ and H₂/Air rotating detonation engine (RDE) with conic aerospike nozzle has been estimated using three-dimensional numerical simulation with detailed chemical reaction model. The present paper provides the evaluation of the specific impulse (I_sp), pressure gain and the thrust coefficient for different micro-nozzle stagnation pressures and for two configurations of conic aerospike nozzle, open and choked aerospike. The simulations show that regardless of the nozzle, increase the micro-nozzles stagnation pressure increases the mass flow rate, the pre-detonation gases pressure and consequently the post-detonation pressure. This gain of pressure in the combustion chamber leads to a higher pressure thrust through the nozzle, improving the I_sp. It was also found that the choked nozzle increases the chamber time-averaged static pressure by 50–60% compared with the open nozzle, inducing higher performance for the same reason explained before.     

Single-ended mid-infrared laser-absorption sensor for time-resolved measurements of water concentration and temperature within the annulus of a rotating detonation engine

A novel single-ended mid-infrared laser-absorption sensor for time-resolved measurements of water mole fraction and temperature was developed and deployed within the annulus of a hydrogen/air-fed rotating detonation engine (RDE). The sensor transmitted two laser beams targeting mid-infrared water transitions through a single optical port on the outer wall of the cylindrical RDE annulus and measured the backscattered radiation from the RDE inner surface using a photodetector for a round-trip path of 1.52?cm. Optimizing the sensor's optical arrangement using numerical ray tracing to minimize interference from optical emission, beam steering, and scattered laser light from window surfaces was essential to sensor performance. Scanned-wavelength-modulation spectroscopy with second-harmonic detection and first-harmonic normalization was implemented to allow for frequency-domain multiplexing of the two lasers and to suppress non-absorbing interference sources such as beam-steering and emission. Tunable diode lasers near 2551 and 2482?nm were modulated at 100 and 122?kHz, respectively, and sinusoidally scanned across the peaks of their respective water transitions at 10?kHz to provide a measurement rate of 20?kHz and detection limit of 0.5% water by mole. Experimentally derived spectroscopic parameters enabled water and temperature sensing with respective uncertainties of 7.3% and 5.3% relative to the measured values. Time-resolved and time-averaged sensor measurements of gas temperature and water vapor mole fraction allow quantitative evaluation of the combustion progress at the measurement location and thus provide a design tool for RDE optimization. Broadly, this single-ended laser sensor should find applications in other combustion systems where optical access is limited.     

Physical characteristics of detonation noise from multi-tube pulse detonation engine

An experimental system of multi-tube pulse detonation engine with single tube to four tubes is set up to study the formation of detonation noise and acoustic ch...     

圆柱空腔内旋转流动中轴对称涡破裂现象的数值模拟

对于上端盖旋转圆柱空腔内旋转流动来说,当流动达到一定条件时,就会出现旋涡破裂现象。目前针对这种现象已经进行了大量的稳态和非稳态实验测量,获得了精确的实验数据。采用数值模拟方法,重现了涡破裂现象及其发生发展过程,同已经发表的相关实验数据进行了比较     

Transverse wave generation mechanism in rotating detonation

Detonation engines are expected to be included in a number of aerospace thrusters in the future. Several types of detonation engines are currently under examination, including the rotating detonation engine (RDE). Although the RDE has been explored experimentally, its rotating detonation propagation mechanism is not well understood. This paper clarifies the detonation mechanism and dynamics of the RDE by 2D and 3D simulation using compressible Euler equations with a full chemical reaction mechanism of H₂/O₂ and H₂/Air, especially from the triple-point and transverse detonation points of view. A total variation diminishing (TVD) scheme is used for the mixture of H₂/Air, and an advection upwind splitting method difference vector (AUSMDV) scheme is used for the mixture of H₂/O₂. The use of an AUSMDV scheme provides a much clearer detonation structure than does the TVD scheme. We focus on the complex interaction mechanism of the detonation front and burned mixture gases. We found out that at this interaction point, an unreacted gas pocket appears and ignites periodically to generate transverse waves at the detonation front and maintain detonation propagation.     

水平滚筒内二元颗粒体系径向分离模式的数值模拟研究

发展了考虑法向接触力、切向接触力和力矩、以及滚动摩擦力矩的三维三方程线性弹性-阻尼离散单元模型及计算程序，对薄滚筒内二元S型颗粒体系进行了数值模拟，发现采用本文的数学模型可以准确地预测出滚筒内二元S型颗粒流的分层现象.分析了影响滚筒内颗粒分层的因素，讨论了滚筒转速、颗粒装载率等参数对分层的影响，当转速较高时，滚筒内形成大颗粒在外、小颗粒在内、具有圆形界面的月亮模式，当转速较低时形成具有波浪形界面的花瓣模式，并且随着滚筒转速的逐渐降低，花瓣的数量逐渐增加，数值模拟结果与实验完全符合.模拟还得到了花瓣模式的形     

水平滚筒内二元颗粒体系径向分离模式的数值模拟研究

发展了考虑法向接触力、切向接触力和力矩、以及滚动摩擦力矩的三维三方程线性弹性-阻尼离散单元模型及计算程序,对薄滚筒内二元S型颗粒体系进行了数值模拟,发现采用本文的数学模型可以准确地预测出滚筒内二元S型颗粒流的分层现象.分析了影响滚筒内颗粒分层的因素,讨论了滚筒转速、颗粒装载率等参数对分层的影响,当转速较高时,滚筒内形成大颗粒在外、小颗粒在内、具有圆形界面的月亮模式,当转速较低时形成具有波浪形界面的花瓣模式,并且随着滚筒转速的逐渐降低,花瓣的数量逐渐增加,数值模拟结果与实验完全符合.模拟还得到了花瓣模式的形 关键词： 分层 滚筒 模式形成 离散单元方法     

Characteristics of pulse detonation engine versus ramjet characteristics

We discuss the method of comparing pulse detonation engines (PDE) and engines with combustion in subsonic flow (ramjet) by means of their specific impulse used by the “Center of Pulse-Detonation Combustion” (CPDC). We demonstrate that the method used by CPDC to calculate the performance of PDE overstates the value of specific impulse relative to its actual value by a factor of at least two. In contrast, the values of specific impulse for ramjet are understated. As a result, the specific impulse of PDE significantly exceeds that of ramjet or is close to it. We investigate these misleading conclusions, and demonstrate their complete failure.