Detonation is an important way to achieve pressurized combustion. Benefiting from the thermal efficiency gain of detonation cycle， the propulsion device based on detonation has significant theoretical performance advantages， which is expected to promote the leaping development of aerospace power technology. On the basis of summarizing the development of classical detonation theory and basic scientific issues related to detonation propulsion， this paper further introduces the research progress of low-order model establishing method of rotating detonation engine， as well as the performance analysis of turbo and ramjet rotating detonation engine， and summarizes the performance analysis and key problem research progress of oblique detonation engine. Finally， some research prospects on detonation engines are discussed based on the current research progresses.

To deal with the difficulty of obtaining accurate mathematic model for turbofan engines and the lacking of guarantee for dynamic performance of a general model reference adaptive control， a multivariable model reference adaptive control method containing an improved reference model is proposed. The multivariable adaptive controller is developed with state feedback for state tracking and the tracking error feedback is introduced into the reference model， which presents apparent restraints upon the overshoot of the tracking curve under large adaptation gains and improves dynamic response. In application to the nonlinear component-level-model of a certain turbofan engine， the control system achieves satisfactory quality of less-than-2s settling time and zero steady-state error.

In order to speed up the development of electric propulsion, this paper compares the development process and application status of electric propulsion technology both in China and abroad. On this basis, for the deep space exploration, commercial aerospace, gravity field measurement, gravitational wave detection and other space missions in China, the thruster should have high specific impulse, long life, wide adjustment range, low cost and high precision. It is proposed that electric propulsion technology should focus on miniaturized ion thrusters, high-power Hall thrusters, pulsed plasma thrusters and drag-free thrusters.

Marine gas turbines dominate the power of large and medium-sized surface warships， which is an important symbol of naval modernization. This paper reviewed the development history of marine gas turbines in the world， and introduced the development of marine gas turbines from Aero-Derivative and independent development. The technical characteristics and performance of typical marine gas turbines LM2500 series， MT30 and UGT25000 were analyzed. Then， the development direction of marine gas turbines in terms of improving power and thermal efficiency， improving reliability and maintainability， developing product serialization and pedigree， etc. are prospected. And， the key technologies that need to be developed in the aspects of materials and manufacturing processes， metal additive manufacturing， aerodynamic design， low emissions， cooling and thermal barrier coatings， dual-fuel， intelligent and hybrid power， etc. are put forward.

To develop a test method for the external drag of high bypass ratio civil aviation turbofan engine nacelle， the NACA-1 series asymmetrical nacelle tests were improved， and the test research of nacelle external drag has been finished in the 2.4m transonic wind tunnel which obtained the external drag of a certain type of high bypass ratio civil aviation turbofan engine nacelle， in the range of Mach number from 0.388 to 0.86. Finally， numerical simulation and experimental results were compared to verify the feasibility of the test method. The results show that the attack angle of the improved test scheme is extended from 0° of the original scheme to 0~4°，the variation of the test data of external drag has a good regularity when Mach number and mass flow coefficient change， the test results agree with the simulation results. The test method could be applied to the measurement of the external drag of high bypass ratio civil aviation turbofan engine nacelle.

In order to solve the problem of the optimization of interfacial gap for submerged nozzle insert throat of solid rocket motor upon thermal loading and aerodynamic pressure， the simulation was performed to study the thermo-structural response of a submerged nozzle. By means of fluid simulation software， the steady flow field of nozzle flow was obtained， and the flow parameters of high temperature gas and convective heat coefficients of nozzle wall were obtained. Meanwhile， based on three-dimensional finite element software with the subroutine of non-uniform pressure and non-uniform heat transfer coefficients， the thermo-structural response of the submerged nozzle was solved， when the value of interfacial gap for throat insert is 0mm， 0.05mm， 0.10mm， 0.15mm， 0.20mm， respectively. The distribution of temperature field and stress field were obtained for the submerged nozzle. The results show that both the hoop compressive stress and tensile stress of throat insert increase at first and then decrease with the increase of time. Secondly， with the increase of interfacial gap， the tensile stress of throat insert decreases at first and then increases， while the compressive stress first increases and then decreases. Thirdly， according to the criterion of interface closure and the minimum hoop stress， the relative optimal value of interfacial gap for throat insert is obtained. The relative optimal value of front gap is 0.1mm， and the relative optimal value of rear gap is 0.05mm. The analyses of gap design and join interface can be conducted by the present numerical method， and the margin about thermal protection and structural strength of the nozzle also can be determined， too.

In order to acquire a better understanding of atomization characteristics under oscillating backpressure conditions, as to gas liquid shear coaxial injector, impinging jet injector and swirl injector that are widely used in liquid rocket engines, the present study reviews recent advances of atomization characteristics of a single liquid jet, gas liquid shear jet, impinging jets, and swirling flow under oscillating backpressure conditions. The main action mechanism of backpressure oscillations on atomization is summarized. Also some problems that existed in previous studies and the key technology that needs breaking through are interpreted. By the review it is concluded that oscillating backpressure affects the atomization field mainly by two ways. One way is to change pressure drop to influence injection thus to affect the atomization process. The other way is to affect atomization directly by oscillating gas field. Plenty of work still needs to be done to study atomization characteristics under oscillating backpressure conditions while some technical difficulties below should be overcome. As for the experiments, backpressure capsule that can generate high-amplitude and high-frequency oscillating backpressure is demanded. At the same time, the disturbance to the atomization field should be at the least. Advanced optical diagnostic apparatus is demanded to acquire atomization field information in the backpressure capsule. As for the simulation, high-fidelity numerical simulation of atomization should be carried out. Also the generation, development, and propagation of pressure wave should be studied. The interaction between oscillating backpressure and atomization field should be investigated based on the two points mentioned above.

The development status of variable cycle turbofan-ramjet (VCTR) engines at home and abroad was summarized, and the operating principles, advantages and disadvantages of two typical configurations of VCTR engines, with or without energy transfer between the turbine engine and the ramjet flowpaths, were comparatively analyzed. The key technologies of VCTR engines are defined and discussed, including engine performance design and simulation technology, high-speed and wide-range fan design technology, after/ram burner (hyperburner) design technology, thermal management system (TMS) design technology, and mode transition design technology. Based on the domestic need and related research status for the VCTR engines, the following issues are proposed to be further investigated: VCTR engine performance design and simulation tool, multi-design points and multi-disciplinary coupling design methodology, TMS design and modeling, and design and test of critical components.

The oblique jet impinging on the wall has a wide range of applications in the field of liquid film cooling in the combustion and atomization of spark injector. In order to study the basic shape， liquid sheet size and the boundary of the liquid sheet formed by the oblique jet impinging on the wall， theoretical modeling studies were carried out. By establishing conservation equations at the boundary of the liquid sheet and the expression of the thickness and velocity distribution of the liquid sheet on the wall， theoretical method for solving the shape and boundary of the liquid sheet formed by the oblique jet impinging on the wall was established. The comparison between the calculation results of the model and the experimental results in the literature shows that the established model can accurately reflect the basic shape of the liquid sheet on the wall. The calculation results show that the liquid sheet spreading area increases with the jet velocity and the jet diameter. When the angle between the jet and the wall increases， the change in the flow distribution causes the liquid sheet length to decrease and the width to increase. The model calculation results can qualitatively reflect that the liquid film boundary becomes smaller with increasing contact angle.

Aiming at the problems in the reliability assessment of liquid rocket engine， such as insufficient utilization of test data， strong subjectivity， and insufficient correlation with design improvements， a two-line assessment method on reliability was presented to improve the accuracy of reliability assessment results under limited test information. In addition， the weak links of engine could be identified from the perspective of risk control. This method used two main lines for reliability assessment. Firstly， taking the engine component test data as input， a reliability model was established based on the engine working sequence and engine failure mode， and the probability of component failure was brought into the reliability model to obtain the prior information of engine reliability using uncertainty propagation method. Then， the engine test data was used to obtain the sample information of engine whose life times was considered to follow a Weibull distribution. Subsequently， the Bayesian method was used for information fusion to calculate the posterior distribution of engine reliability. This method expands the amount of data for reliability assessment， identifies the key component failures of the engine from a risk control perspective， and improves the engineering application value of reliability assessment.

High power nuclear electric propulsion (NEP) highly integrates space nuclear reactor and high power electric propulsion technologies, with the advantages of high power density, very high specific impulse and high thrust, which can be applied for large interplanetary missions, such as orbital transfer task of super large spacecraft, robotic interplanetary missions and manned Mars exploration, greatly enhancing human ability in deep space exploration. This paper presents the principle and system configuration of high power NEP system, and summarizes the progress of several critical techniques. Moreover, both research process and recent development in worldwide are also outlined.

The advantages of underwater fuel cell propulsion are high energy conversion efficiency and energy density， low noise and no gas emissions， thereby improving the performance of unmanned undersea vehicles （UUVs） in terms of range， depth and stealth. As such， underwater fuel cell propulsion is a very promising candidate for future UUVs. This paper introduces the system components and working principle of the underwater fuel cell propulsion system. The recent progress on key technologies is reviewed， including fuel cell-powered UUVs， hydrogen-oxygen fuel cell and energy-dense reactant storage. Future research work in this field is also discussed. In terms of hydrogen-oxygen fuel cells， pure oxygen supply and closed-cycle operation would bring water removal and corrosive issues， which should be carefully dealt with. In terms of hydrogen and oxygen sources， high energy density formats include water-reactive aluminum and diesel reforming for hydrogen acquisition， lithium perchlorate for oxygen acquisition， which deserve increasing research.

The classifications and characteristics of deep-sea underwater vehicles are introduced， then the way in which the propulsive force of each deep-sea underwater vehicle is generated is analyzed respectively， and the main types and characteristics of the thrusters of deep-sea underwater vehicles are summarized. Altogether， the configuration and structural characteristics of 5 typical types of deep-sea underwater vehicle thrusters are introduced. And according to the characteristics of failure mode of the thruster’s failure， six types of common failure modes， incentives and common diagnostic methods of deep-sea underwater vehicle thrusters are presented. Qualitative analysis of diagnostic methods， analytical model diagnosis methods and signal processing diagnosis methods are used to summarize current research progress of fault diagnosis technology for deep-sea underwater vehicle thrusters， and the suggestions for the development of deep-sea underwater vehicles and their failure diagnosis technologies are put forward.

In order to study the effects of wall temperature on the flow field structure and flow parameters of oblique detonation engine inlet in laminar， transition and turbulent states， an oblique detonation engine inlet with curved compression section at Mach 10 was selected as the research object to conduct numerical simulation， and the changes of shock induced separation zone and thermal boundary layer near the inlet wall were deeply discussed. The numerical simulation results show that the wall cooling can restrain the phenomenon of relaminarization in the circular transition section of the inlet shoulder， and the relaminarization is the most serious under the condition of adiabatic wall. The increase of wall temperature is beneficial to delay the flow transition， and it also leads to the increase of separation area size and the moving forward of the main body of separation zone in transition and turbulent state. The transition in the inlet channel is separation induced transition. The transition position is mainly affected by the position of separation point. The overall performance is that the transition position gradually moves forward with the increase of wall temperature. With the increase of wall temperature， the thickness of thermal boundary layer on the top plate side of the inlet exit gradually thickens. In transition state， the thermal boundary layer thickness changes by up to 5%. The peak temperature on the top plate side of the inlet exit increases with the increase of wall temperature， and the peak position is gradually close to the wall. When the wall temperature is the same， the heat flux and thermal boundary layer thickness is smaller in the laminar flow state. In the transition and turbulent state， the thermal boundary layer on the top side of the inlet exit is thicker， which is about three times that of laminar flow， and the difference of thermal boundary layer thickness between transition and turbulent state can reach 2%.

Numerical simulations and experiments of a Ma 9 oblique detonation engine were conducted to study the combustion mechanism of oblique detonation engines. Firstly， a full-scale engine model with a length of 2.8m was designed. The engine inlet is a two-stage compression inlet composed of two 15°-inclined ramps. The hydrogen is pre-injected into the main flow at the leading front of the inlet by three strut-injectors. Secondly， the mixing process at the inlet and the combustion process in the combustor were numerically simulated. In the numerical simulations， the governing equations are Reynolds average Navier-Stokes equations with SST k- ω turbulence model and 9-species and 19-recations detailed chemical reaction kinetics. The numerical results show that the mixing process of hydrogen along the inlet is good. The stable oblique detonation waves and normal detonation waves are obtained in the combustor. Finally， the experiments under Ma 9 flight conditions were conducted in the shock tunnel. The stable flow fields of oblique detonation engine were established in the duration of 50ms test time of the shock tunnel. The experimental results are in good agreements with numerical results which means that stable oblique detonation waves were successfully obtained in the shock tunnel experiments. This research results demonstrate the technical feasibility of oblique detonation engines.

In order to analyze the law of oscillating combustion induced by spheres with different diameters and to reveal the inherent effect of sphere size on the oscillation phenomenon， numerical simulations are carried out to investigate the oscillating combustion phenomenon in a H ₂/air premixed gas mixture induced by spheres with different diameters， by solving the two-dimensional axisymmetric Euler equations along with a detailed combustion mechanism. Results show that as the sphere diameter increases， the frequency of oscillating combustion does not decrease continuously， but with two abrupt drops， which implies that there exist three modes in the high-speed sphere-induced oscillating combustion phenomenon， namely the superhigh-frequency mode， the high-frequency mode， and the low-frequency mode. During the transition of two modes， there exists a metastable oscillating state of double-frequency coupling before the oscillation reaches its stable state. Moreover， the appearance of these three different modes is affected by different oscillation mechanisms， and the double-frequency coupling phenomenon during the transition of two modes is resulted from the competition of two mechanisms.

Numerical techniques for multi-species chemical reacting flows are applied to study the macrostructures of oblique detonation waves in confined space. The key features， the evolution of the overall wave structure， and the transitional criteria are analysed. The results show that with the increase of wedge angle， four structures appear in turn： shock-induced combustion， oblique detonation double regular reflection， Mach reflection in reflux zone， and wedge combustion. The simulations reveal three critical conditions of the compression angle for the maintenance of stable wave structures， i.e.， lower-critical， sub-critical， and upper critical conditions. In addition， the transition between regular and Mach reflections occurs during the evolution process of the macro wave structure.

An Accurate thermal prediction of main shaft roller bearings for aero-engine is key to ensure stable operation of engine lubrication system and the whole engine. Considering that traditional thermal network method is less accurate， a finite-element based thermal network method is proposed. The bearing is divided into a finite number of elements by mesh， and each element is provided with a temperature node representing lumped parameters of the element. Then thermal resistance relationship is established among adjacent nodes to form a thermal network. By solving the large sparse matrix linear equations， the temperature distribution in the cross section of roller bearing could be obtained. This method could be combined with the distributed bearing heat generation theory to achieve accurate local heat loading and precise temperature distribution analysis. By comparing with experimental results， it is found that the error of predicted centerline temperature of outer ring surface in the bearing is less than 13%.

Vacuum plume in real environment is uncertain whereas the numerical simulation is always deterministic, thus the influence of uncertainty on the vacuum plume is required to be elaborately studied. In this paper, the direct simulation Monte Carlo (DSMC) method was used to simulate the plume flowfield with various input uncertainties. Meanwhile, the probabilistic collocation method with sparse grid technique was used to describe the input uncertainties stemming from the incoming flow, wall condition and model parameters. The mean, variance and uncertainty of the output parameters, and the propagation of uncertainty were also calculated by the probabilistic collocation method. The corresponding results indicate that the uncertainties in flowfield sharply increase along the streamline when the flow speed reaches its maximum, and then decrease when close to the sound speed line. The influence of input uncertainty in wall temperature dominates in the downstream area of the sound speed line. Note that, the increment of uncertainty in the pressure is the most significant, which approximately equals to 2.1 times the input uncertainty (3.54%). In addition, the temperature jump uncertainty is almost constant, which is about 0.8 times the input uncertainty. This may be caused by the input uncertainty in the wall temperature that bounds the temperature jump uncertainty. As a consequence, the uncertainty in the heat flux (5.54%) is slightly smaller than that in the normal stress (6.25%), meanwhile the uncertainty in the shear stress (5.07%) is the smallest one. Finally, the Sobol’ global sensitivity analysis shows that the input velocity uncertainty and pressure uncertainty in the throat contribute the most to the uncertainty in aerodynamic characteristics, and considerably exceed the influence of the input uncertainties in the wall temperature and in the reference diameter of simulated particles.

Shock-induced combustion of a blunt projectile is a basic problem in detonation research. Numerical simulations were conducted to study the shock-induced combustion phenomenon of a stoichiometric H ₂/Air mixture at the flow Mach numbers of 4.79 and 6.46. The block-structured adaptive mesh refinement program AMROC based on the finite volume method was adopted to solve the axisymmetric Euler equations with chemical reaction source terms， and the influence of some important factors such as the form of the MUSCL reconstruction， the slope limiter types， and the chemical reaction mechanisms were investigated. The results show that， based on the mesh adaption flag parameters， the program can realize adaptive mesh refinement efficiently. The comparisons with experimental results show that the accuracy of the unsteady shock-induced combustion simulation depends not only on the chemical reaction mechanism， but also on the form of the limiter. Adopting two different forms of the MUSCL reconstruction format acquires almost the same oscillating frequencies， which is 1.17% and 0.97% different from the result obtained in the experiment， respectively. Numerical study is conducted for comparing the classic Jachimowski mechanism with several newly developed pressure-dependent hydrogen/oxygen reaction mechanisms. It is shown that， in the unsteady shock-induced combustion case at Ma=4.79， the classic Jachimowski mechanism is still the most suitable mechanism to obtain the closest oscillating frequency to the experimental result. While in the steady shock-induced combustion case at Ma=6.46， all of the given mechanisms can give results that are in good agreements with the experiment.