Aerospace Mechatronics and Control Technology

The drag and heat flux reduction have become an increasing crucial characteristic for the hypersonic vehicles. Numerical simulations of the hypersonic flow around the blunt body with aerospike are investigated for the aerospike’s changing L/D ratios with freedom Mach number of 5.75. The effect of the characteristics of the flow field and the aerospike’s L/D ratios on the drag coefficient of model and the surface pressure coefficient of the nose cone are analysed using a two-dimensional axisymmetric Reynolds-averaged Navier–Stokes (RANS) equations coupled with the two equation k-ω shear stress transport (SST) turbulence model. The numerical simulation results show that the installation of the aerospike changes the flow field structure of the blunt body vehicle head, which reduces the drag of the blunt body vehicle significantly in hypersonic flight. The size of the recirculation region depends on the aerospike’s L/D ratios. Among the different aerospike configurations studied, the drag reduction effect of aerospike when the L/D = 1 is the best, and the maximum value of the drag reduction efficiency is 36.4%. Within the scope of this study, as the L/D ratio of the aerospike increases, the drag reduction efficiency increases gradually.

In order to solve the terminal guidance problem of the hypersonic vehicle, a multi-constrained guidance law with position, impact angle and impact velocity was studied. Firstly, a finite time sliding mode control method based on truncation function for second order nonlinear system was proposed, and then a reasonable truncation function was designed according to the dynamic equation of the terminal guidance phase of the vehicle to form the sliding mode control function of the longitudinal and transverse planes. Further, the auxiliary control was obtained and the guidance command was generated. Under the action of the control quantities, the system was kept on the sliding mode surface all the time, that is, the global sliding mode can be realized, and the constraints of position and impact angle can be realized. The guidance parameter was solved iteratively according to the redundant degrees of freedom in the transverse plane to ensure the constraint of impact velocity. The simulation verified that the guidance method can meet the requirements of terminal multiple constraints of hypersonic vehicle.

Air rudder is the actuating mechanism of the boost-glide aircraft, which is used to control the flight attitude of the aircraft. Before the flight test, the time domain and frequency domain identification should be carried out by measuring the performance indexes of the air rudder, which is used for the 6-DOF simulation of the aircraft. This paper designs the test platform and test scheme of the air rudder of the boost glide aircraft, and tests the performance of the air rudder. The time domain and frequency domain identification methods are used to identify the four air rudder models of the boost-glide aircraft in this paper, and the verification scheme of cross-comparison between the time domain identification model and the frequency domain identification model is proposed. The test results show that the time domain identification method has low precision and can only represent the steady-state response of the air rudder. The frequency domain identification method has high accuracy, which is suitable for characterizing the change process of the air rudder from the initial response to the stable state, and the identification results can be applied well in the attitude control system of the boost-glide aircraft.

In order to study the flow characteristics and the heat flux of cavity with length-to-depth ratio of 6 on the surfaces of the hypersonic reentry aerospace vehicles in rarefied gas flow, Direct Simulation Monte Carlo method (DSMC) with the adaptive grid is used. The free stream at a Mach Number of 8, an altitude of 60 km, and the angles of attack (AOA) ranging from 0° to 60°.The results show that the closed cavity changes its type when AOA changes and is back to closed type when AOA is 60°. Increasing AOA sometimes does not help free-stream to get into the cavity because the stream strikes the plate on the front edge of the cavity, and form a shock wave that changes the direction of the free-stream flow. The low-speed and high pressure region inside the cavity extends to the upper zone of the cavity due to the shock wave, the speed is getting smaller and the pressure is getting higher. When AOA is 40° and 50°, the heat flux of three surfaces of the cavity is higher than other situation.

In this paper, firstly the reasons of aero-engine bearing skid were analyzed from these aspects of fault phenomenon, skid mechanism, manufacturing process, practical use, trial run, measurement and inspection by fishbone diagram. In order to ensure the economy and airworthiness of aero-engine manufacturing, the use and repair of bearing with skid damage were discussed by means of information statistics and trial run experiments. Then the effect of bearing skid on the engine safety was analyzed. Based on the above research, the possible causes and preventive measures for bearing skid are expounded. And the repair method, experimental content and inspection standard of bearing with skid are presented in detail. The research has important guiding significance for aero-engine bearing to master its working state, prevent its fault and guarantee the aviation safety.

Aero-engine blade tip clearance (BTC) is one of the most important parameters that have decisive influence on engine performance. Thus, it is a typical inspection indicator during engine assembly. With the higher demand for consistency of engine performance, the measurability of BTC is required to be much more accurate and dynamically. However, the traditional BTC detection method in the assembly process is restricted to lower accuracy, inability of complex structures and lack of dynamic measuring capability. This paper introduces a visual BTC measuring system based on structured light, which can directly measure the clearance between rotor and stator casing by observing its position from the engine principle axial direction. A special laser generator is designed to project structured light to the surfaces of blade and casing. Simultaneously, the light is captured by the industrial camera. Compared to the traditional methods, the BTC in whole circumferential phase can be measured rather than single phase. Meanwhile, the proposed method is not influenced by casing deformation and installation errors. Finally, the feasibility and capability of the system is verified by simulation experiment.

In modern aero-engine assembly production, the assembly deviation analysis method plays an important role in product quality and efficiency. Various mathematical model based analysis methods are studied to realize assembly deviation prediction. However, most established analysis methods require abstraction and simplification of geometric deviations, which results in the inability to accurately characterizing the surface deviation features and analyzing the impact of deviation characteristics on the assembly. Therefore, a Skin Model Shapes based assembly deviation analysis method is proposed to enhance the accuracy and efficiency of deviation identification and assembly. This method enables the analysis of the assembly deviation caused by the surface deviation feature. The feature surface is represented by Skin Model Shapes, which is generated from surface measurement data. An assembly experiment of high pressure turbine components is performed to illustrate the procedure of this method, and the effectiveness of the method is verified by comparing the experimental and simulation results.

Considering the ground-to-flight difference in solid rocket motor and uncertainties on motor performance parameters at the shutdown point, this paper proposes a method for online identification and prediction of solid rocket motor performance. The calculation model of the nozzle throat diameter and the propellant burning rate are established based on the analysis of internal correlation between motor parameters. Then by means of the real-time measured data from the missile-borne sensor, throat diameter and burning rate can be calculated. After the model parameters of throat diameter and burning rate were identified by the recursive least square algorithm, the shutdown time and thrust-time characteristics can be predicted based on identification results and prediction model. The simulation results show that this method has high identification and prediction accuracy, and could provide a feasible solution for online identification and prediction of the performance parameters of solid rocket motor.

Due to the error of target indication, the existing laser beam riding guidance weapon system is difficult to intercept the long-range target accurately. In order to solve this problem, a multi-missile coverage interception strategy is proposed in this paper: the total number of missiles determined is decomposed into a finite batch, and the interception probability is improved by adjusting the position of the same group of missiles in the interception plane. In addition, this paper proposes an Improved Grey Wolf Optimizer (IGWO). Based on the traditional Grey Wolf Optimizer, this algorithm generates the initial population uniformly by Hammersley sequence, and adopts a nonlinear convergence factor adjustment method, which can quickly and effectively optimize the location of the missiles. The simulation results show that the improved grey wolf optimizer improves the search accuracy, stability and convergence speed greatly. The multi-missile coverage interception strategy based on the improved grey wolf optimizer can effectively improve the interception accuracy.

The mirror subassembly is the key component of the space optic remote sensor and the mirror shape is always required to reach high accuracy due to the crucial importance to the system’s imaging quality. However, the mirror shape is usually influenced by gravity, temperature, forced displacement and launch vibrations to distort after leaving earth. The support structure of the subassembly plays an important role to minimize the impact of above factors while the conflict of the mirror’s static accuracy and the subassembly’s dynamic strength needs to be studied. Therefore, based on a Φ 260 mm-aperture mirror subassembly, a whole set of flexible support structures was researched. Combining with the conventional ring-type lateral support structure, a discrete-type flexible mounting scheme was proposed. Materials of the mirror and cell, the micro-stress mounting strategy and parameters of the flexure hinges were established aiming at maintaining high static precision of the mirror under multi-conditions, and the resilient connector was optimized by stress deconcentration and viscoelastic materials damping to obtain adequate dynamic strength margin. Finally, the comprehensive performance of the subassembly was verified eligible by finite element analysis that the mirror surface error RMS was superior to 0.002 λ under objective static conditions with enough safety margin in the harsh vibrations. In conclusion, the flexible support structure brought forward by this article is feasible which has general compatibility for other small-size space mirror.

In order to explore the effect of the overall vortex flow distortion on the performance of the low-speed axial compressor, this paper carried out the compressor characteristic experiment under the overall vortex flow distortion generator based on the overall vortex flow distortion generator, and measured the performance of the swirl flow distortion generator. The generated swirl field obtains the pressure ratio and efficiency characteristics of the compressor under the conditions of uniform intake and swirl intake. The results show that the overall vortex flow distortion increases the pressure ratio of the compressor, reduces the efficiency of the compressor, and at the same time causes the stability margin of the compressor to decrease.

For image-guided missile, it is important to localize the target and its surrounding area using the images taken by seeker. In this paper, we propose a target localization method that combines on-board IMU data and image information. This method does not require equipment such as GPS or optoelectronic platform, nor does it need prior knowledge or templates of the target. Firstly, a selection strategy is proposed to determine keyframes, which are used to solve the pose of the seeker camera and the 3D coordinate of the target in the camera coordinate system. Then nonlinear optimization is used to reduce the reprojection error and obtain the optimal solution of the seeker camera pose and the 3D coordinate of the target. Finally, the scale uncertainty of the seeker monocular camera is solved by combining the on-board IMU data with images to obtain the coordinate of the target in the shooter coordinate system. The proposed method is validated in the flight data of a certain type of image-guided missile, and a high-precision localization result is obtained at a long missile-target distance.

The high dynamic characteristics of agile micro-nano satellites will introduce noncommutativity error in the attitude updating process. Therefore, the attitude updating algorithm of the traditional micro-nano satellite is no longer applicable. This paper proposes a modified attitude updating scheme applied to agile satellites. The method adopts the coning motion as the research scene, in which the influence of noncommutativity error is the worst. Firstly, traditional attitude updating algorithm is presented, and based on the shortcoming of the algorithm, the rotation vector algorithm is introduced. Then, combined with the agile micro-nano satellite's characteristics, the upper bound and lower bound of the algorithm's sampling number are given. Besides, referring to the least-squares algorithm, we carry out angular velocity fitting and calculate the required equal interval angular velocity. Finally, the rotation vector compensation algorithm is derived to gain the updating attitude. The proposed algorithm does not need equal interval angular velocity sampling and can achieve higher estimation accuracy in a dynamic environment than the traditional algorithms. The simulation results validate the effectiveness of the proposed scheme.

Autonomous system design has received extensive attention for orbit and attitude determination, since the traditional ground station-based orbit determination is difficult to meet the multi-satellite needs. With a combination of limited sensors, this article presents an autonomous navigation-attitude determination system for low-earth-orbit micro-nano satellites. By adding the area-to-mass ratio to state vector, the atmospheric resistance is effectively considered. By introducing an infrared earth sensor, the limitations of the orbit type and eclipse period are compensated, with the consideration of computational burden. Simulation based on ZDPS-2 satellites show the estimation accuracy of this paper is improved by 23% compared with the magnetometer/sun sensor combination, reaching 1.08 km, 1.16 m/s (RMS); whereas for an equatorial orbit, estimation accuracy remains 1.11 km (RMS). While completing navigation estimation, the system attitude pointing accuracy reaches 0.347° (RMS), which meets the basic mission requirements of micro-nano satellites.