Aerospace Systems最新文献

An axial single-stage high-speed test rig is numerically studied in this paper with half-annulus URANS simulations to describe the flow characteristics at the near stall condition. Wavelet analysis is applied to demonstrate the time–frequency characteristics of the near-tip pressure signals captured by the numerical probes at different circumferential and axial positions. The detailed tip flow fields and wavelet transform results are combined to depict the generation and propagation of the spike-type stall inception. According to the wavelet spectrum, characteristic frequencies correspond to the temporal and spatial features of the rotating stall, such as the fluctuation of the shock wave, self-oscillation and propagation of tip leakage vortex et al. Consequently, the detection of typical spike stall inception can be significantly brought forward by identifying the crucial rotating disturbance and its development for the onset of stall inception. Then, the specific tip flow fields are also discussed to reveal the flow mechanism of stall inception evolution, including the leading edge spillage and the trailing edge backflow. Further investigation shows that the stall inception with smooth casing corresponds to the radial separation vortex caused by the tip leading edge spillage, which continues to develop and propagate in the circumferential direction and finally induces the stall.

This paper investigates the circulation, vorticity, and oblique shock waves generated by the airfoil with a spoiler. In this regard, the effects of the geometric parameters of the airfoils, such as the attack angle, shock wave angle, and Mach number etc., are studied first. We have introduced the conformal mapping technique, which has helped transform the flow past a circular cylinder to an airfoil shape. Based on the Kutta conditions, we have solved the unsteady flow in the region of an airfoil. The results show that the flow patterns form a strong vortex after the spoiler when the attack angle of the airfoil changes from 0° to 10° at a spoiler angle of 85°. Furthermore, we have used the shock wave expansion method, and the deflection angle has been found to increase with increasing the shock wave angle for a certain Mach number. A higher shock wave angle is obtained for a lower Mach number at any deflection angle. In addition, it is also found that the post Mach number is decreased with the increasing shock wave angle for fixed pre Mach number.

The turbulent transonic two-dimensional airflow in 9°-bent channels is studied numerically on the basis of the Reynolds-averaged Navier–Stokes equations. The flow is supersonic at the entrance of channels and subsonic at the exit. Numerical solutions reveal non-uniqueness of flow regimes in certain ranges of boundary conditions. The location of a formed shock wave exhibits hysteresis with changes in the inflow Mach number M_∞, or the angle of attack, or pressure given at the exit p_exit. The existence of hysteresis is caused by an interaction of the shock wave with an expansion flow region over the convex wall of channel. Shock wave behavior under forced oscillations of the Mach number M_∞ or pressure p_exit is discussed. Dependencies of hysteresis and non-unique regimes on the amplitude and period of the oscillations of M_∞, p_exit are studied. It is shown that hysteresis in a long channel is essentially wider than that in a short one.

This paper presents the development and analysis of a spacecraft formation flying architecture. The desired state of each spacecraft is maintained using a model predictive control-based control framework that is based on the Hill–Clohessy–Wiltshire equations and a polytope boundary constraint as a switching surface. This framework can be used to maintain the desired cluster formation while also guaranteeing internal cluster flight. The polytope boundaries are designed, such that no two agents have overlapping regions, allowing the vehicles to execute avoidance strategies without continually maintaining the trajectories of other agents. The model predictive control framework combined with the convex polytope boundary enables a scalable method that can support clusters of satellites to coordinate to safely achieve mission objectives while minimizing fuel usage. As part of the implementation of this control scheme, the authors created two spacecraft formation flying control approaches. The first approach uses fewer, large maneuvers to control a spacecraft to the center of a keep-in-volume. The second approach allows the spacecraft to perform many small maneuvers to stay just inside the boundary of the keep-in-volume. This paper compares the fuel cost savings of these two approaches. The results presented in this paper demonstrate that the first approach produces the lower total fuel usage, but if a lower amount of fuel per maneuver is required, then the second approach should be used. This work also compares the computation requirements and fuel usage for (hbox {L}_1), (hbox {L}_2), and (hbox {L}_infty ) norms formulations of the framework, the (hbox {L}_1) and (hbox {L}_2) norms require the least amount of fuel usage, while the (hbox {L}_2) requires the least amount of computation time.

Recent development in space mission demands safer and more cost-effective space missions. Hybrid rocket engine technological advancements have prolonged a critical stage in their development and it is the better option for such space missions, as it has a lot of advantages over the solid rocket motor and liquid rocket engine. It is simple in design, has high thrust density, low weight, and is safer than a liquid rocket engine. It has restarted capability, safe, low explosion risk, and high specific impulse than a solid rocket motor. This paper shows the numerical analysis of a hybrid rocket engine. The paper highlights the initial boundary conditions in the analysis of a 300-N hybrid rocket engine. The process started with a chemical kinematic examination of engine-compatible fuels and oxidizers. This investigation provided the fundamental parameters required for the design and subsequent dimensioning of a hybrid rocket engine. It also produced a three-dimensional design model, performed numerical analysis using ANSYS software, and validated the findings using existing literature. Using the k–(varepsilon) turbulence model and transient solver on 8 mm port diameter for analyzing. The computational fluid dynamics model offered the qualities of a real hybrid rocket engine and it will be helpful to researchers and the scientific community in the future.

The three degrees of freedom spacecraft attitude simulator is of vital importance in verifying spacecraft control strategies and many other space techniques. It requires accurate knowledge of simulator inertia parameters which can be identified by a variety of estimation methods under appropriate excitation situation. However, constraints on the rotation range, angular velocity, and torque may lead to a bad parameter estimation performance and cause security problem in excitation process. A new adaptive reorientation controller is proposed in this paper to solve these problems. By deriving the expression of parameter estimation error and analyzing the ill-conditioned problem resulted from the attitude constraint, a preconditioned adaptive parameter estimation law is designed and then combined with a new proposed reorientation controller, such that the errors of parameter identification and reorientation excitation simultaneously converge to zero. And the constraints can also be met. Compared to conventional parameter identification schemes, the proposed controller can simultaneously achieve the closed-loop reorientation excitation for security requirement and the more efficient parameter identification outcome. The effectiveness of the adaptive controller is finally demonstrated by numerical simulations.

In this work, a mixed Kalman/H-infinity filter is designed for the attitude estimation of a low Earth orbit microsatellite and the external disturbance torques acting on it. The state vector will be formed by satellite's attitude along with angular rates and the external disturbances. An improved external disturbance modeled as a random walk acting (slowly varying) around three axis attitude was proposed. This external disturbance is mainly generated by the aerodynamic torque, the residual magnetic moment and the gravity gradient torque. The satellite has only magnetometer on board as the attitude sensor. The proposed algorithm is tested using simulated data for a microsatellite, and the results of this study are tested in different scenarios. The first two scenarios are the cases with and without uncertainty in the satellite’s inertia. The last scenario is extensive Monte Carlo simulations with uniformly distributed initial conditions of the Euler angle and angular rate. The major purpose of this work is to demonstrate that we can estimate external disturbances and attitude dynamic parameters of a satellite using a simple filter that combines the best features of Kalman and ({mathrm{H}}_{infty }) filters. The simulation results show that the attitude RMS error is less than (pm 1) deg (acceptable accuracy). Also, Monte Carlo simulation gives good results of the proposed filter. This latter estimates the attitude with accuracy less than 0.8 deg, the rate order is 1 milli-deg/s and the external disturbances around 1.5 μNm.

This paper describes the main projects for the creation of supersonic administrative aircraft and the features of the weight analysis of supersonic passenger administrative aircraft. The question of the resource and strength of the structure in operating conditions at high temperatures is considered from the point of view of weight analysis. A model is proposed for calculating the mass of the supersonic passenger administrative aircraft fuselage in the first approximation based on regression analysis.

UAVs have been increasingly used in military and commercial applications. The theory of UAV swarm behavir has gradually matured and moved to the real application stage. Fast and accurate recognition of the intentions of UAV swarms become a key part of dealing with coming swarms. This paper proposes a data-driven approach to realize the recognition of the typical intentions of UAV swarm. The UAV swarm’s intention is divided into three basic categories: expansion, free movement, and contraction. The dubins model is introduced to depict and study the dynamic characteristics of the movement of the UAV swarm. Simulation experiments are performed through software to collect data and to verify and refine the proposed data-driven intention recognition approach. Moreover, real flight experiments are conducted to test the feasibility and accuracy of the proposed approach, from which key steps about the neural network building and training for intention recognition have been summarized, and satisfying results in intention recognition with high accuracy and stability during the entire movement of the UAV swarm have been achieved.

The paper considers the smoothing of tabulated curves describing airfoils. Smoothing is required to eliminate airfoil contour distortions that occur during the design of the aircraft wing surface. The problem of ensuring a smooth change in the curvature of the smoothed contour is presented as a problem of minimizing the quadratic function of many variables. To minimize the objective quadratic function, the gradient descent method with a constant step was used. According to the developed technique, with the help of a computer program, the smoothing of the airfoil was carried out. As a result, a rather smooth diagram of the profile curvature was obtained, which confirmed the effectiveness of the developed smoothing technique.