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This paper aims to investigate the problem of attitude control for a horizontal takeoff and horizontal landing reusable launch vehicle.

In this paper, a predefined-time attitude tracking controller is presented for a horizontal takeoff and horizontal landing reusable launch vehicle (HTHLRLV). Firstly, the attitude tracking error dynamics model of the HTHLRLV is developed. Subsequently, a novel sliding mode surface is designed with predefined-time stability. Furthermore, by using the proposed sliding mode surface, a predefined-time controller is derived. To compensate the external disturbances or model uncertainties, a fixed-time disturbance observer is developed, and its convergence time can be defined as a prior control parameter. Finally, the stability of the proposed sliding mode surface and the controller can be proved by the Lyapunov theory.

In contrast to other fixed-time methods, this controller only requires three control parameters, and the convergence time can be predefined instead of being estimated. The simulation results also demonstrate the effectiveness of the proposed controller.

A novel predefined-time attitude tracking controller is developed based on the predefined-time sliding mode surface (SMS) and fixed-time disturbance observer (FxTDO). The convergence time of the system can be selected as a prior control parameter for SMS and FxTDO.

The purpose of this paper is to develop an attitude control strategy for the reusable boosted vehicle with large angle of attack, and to remove the cross coupling among roll, pitch and yaw channels.

The coordinated gain scheduling control strategy consists mainly of two parts. First, initially ignoring dynamic coupling, single channel gain scheduling controller is designed based on linearized models, respectively. Second, with respect to main channel gain scheduling controller, coordinated scheduling controller is used to generate intentionally cross coupling to partly cancel inter‐channel cross coupling of reusable boosted vehicle.

A coordinated gain scheduling control strategy is presented, and no such analytical solution can be found for the reusable boosted vehicle.

The design idea of coordinated gain scheduling strategy is straightforward in physical concepts and has great value for engineering applications.

Coordinated gain scheduling control strategy is novel in that single channel gain scheduling design does not involve small perturbation linearization and coordinated channel is scheduled.

The purpose of this paper is to present a sliding mode attitude controller for reusable launch vehicle (RLV) which is nonlinear, coupling, and includes uncertain parameters and external disturbances.

A smooth second-order nonsingular terminal sliding mode (NTSM) controller is proposed for RLV in reentry phase. First, a NTSM manifold is proposed for finite-time convergence. Then a smooth second sliding mode controller is designed to establish the sliding mode. An observer is utilized to estimate the lumped disturbance and the estimation result is used for feedforward compensation in the controller.

It is mathematically proved that the proposed sliding mode technique makes the attitude tracking errors converge to zero in finite time and the convergence time is estimated. Simulations are made for RLV through the assumption that aerodynamic parameters and atmospheric density are perturbed. Simulation results demonstrate that the proposed control strategy is effective, leading to promising performance and robustness.

By the proposed controller, the second-order sliding mode is established. The attitude tracking error converges to zero in a finite time. Meanwhile, the chattering is alleviated and a smooth control input is obtained.

The purpose of this paper is to address the flaws of traditional methods and fulfil the special fault‐tolerant re‐entry navigation requirements of reusable boost vehicle (RBV).

A kind of improved estimation method based on strong tracking unscented Kalman filter (STUKF) is put forward. According to the fact that the traditional state χ²‐test‐based fault diagnosis method is incompetent to detect the signal point small jerks and slowly varying fault in the measurement, a kind of original fault diagnosis technology based on STUKF is used to check the working states of navigation sensors.

The comparisons with χ²‐test method under typical failure distributions validate the perfect state tracking and fault diagnosis performances of this improved method.

This kind of state estimation and fault diagnosis method could be used in the navigation and guidance systems for many kinds of aeronautical and astronautical vehicles.

A kind of novel strong tracking state estimation filter is used, and a kind of very effective fault diagnosis criterion is put forward for the navigation of RBV.

The capability to predict and evaluate various configurations’ performance during the conceptual design phase using multidisciplinary design analysis and optimization can significantly increase the preliminary design process’s efficiency and reduce design and development costs. This research paper aims to perform multidisciplinary design and optimization for an expendable microsatellite launch vehicle (MSLV) comprising three solid-propellant stages, capable of delivering micro-payloads in the low earth orbit. The methodology’s primary purpose is to increase the conceptual and preliminary design process’s efficiency by reducing both the design and development costs.

Multidiscipline feasible architecture is applied for the multidisciplinary design and optimization of an expendable MSLV at the conceptual level to accommodate interdisciplinary interactions during the optimization process. The multidisciplinary design and optimization framework developed and implemented in this research effort encompasses coupled analysis disciplines of vehicle geometry, mass calculations, aerodynamics, propulsion and trajectory. Nineteen design variables were selected to optimize expendable MSLV to launch a 100 kg satellite at an altitude of 600 km in the low earth orbit. Modern heuristic optimization methods such as genetic algorithm (GA), particle swarm optimization (PSO) and SA are applied and compared to obtain the optimal configurations. The initial population is created by passing the upper and lower bounds of design variables to the optimizer. The optimizer then searches for the best possible combination of design variables to obtain the objective function while satisfying the constraints.

All of the applied heuristic methods were able to optimize the design problem. Optimized design variables from these methods lie within the lower and upper bounds. This research successfully achieves the desired altitude and final injection velocity while satisfying all the constraints. In this research effort, multiple runs of heuristic algorithms reduce the fundamental stochastic error.

The use of multiple heuristics optimization methods such as GA, PSO and SA in the conceptual design phase owing to the exclusivity of their search approach provides a unique opportunity for exploration of the feasible design space and helps in obtaining alternative configurations capable of meeting the mission objectives, which is not possible when using any of the single optimization algorithm.

The optimized configurations can be further used as baseline configurations in the microsatellite launch missions’ conceptual and preliminary design phases.

Satellite launch vehicle design and optimization is a complex multidisciplinary problem, and it is dealt with effectively in the multidisciplinary design and optimization domain. It integrates several interlinked disciplines and gives the optimum result that satisfies these disciplines’ requirements. This research effort provides the multidisciplinary design and optimization-based simulation framework to predict and evaluate various expendable satellite launch vehicle configurations’ performance. This framework significantly increases the conceptual and preliminary design process’s efficiency by reducing design and development costs.

Space tourism has to be regulated as a subset of private spaceflight activities, whereby humans are sent to outer space in a fundamentally private context. In addition to space law, air law would be relevant for addressing private spaceflight, but neither regime has at the international level regulated relevant activities to any appreciable extent. They provide little more than a set of guiding overarching principles. Much of the onus of future regulation will fall on the shoulders of individual states, most notably the United States. In the more distant future, this may result in a special international regime, using elements of both space and air law.

The purpose of this paper is to propose a hybrid optimization approach with high level of solving precision and efficiency for endo-atmospheric ascent trajectory planning of launch vehicles.

Based on the indirect method of optimal control problems, the optimal endo-atmospheric ascent problem with path constraints and final condition constraints is transformed into a Hamiltonian two point boundary value problem (TPBVP). An advanced Gauss pseudo-spectral method is applied to change the Hamiltonian TPBVP into a system of nonlinear algebraic equations, which is solved by a modified Newton method. To guarantee the convergence of the solution, analytical initial guess technology and homotopy technology are also introduced. At last, simulation tests are made.

The hybrid approach for optimal endo-atmospheric ascent trajectory planning has both fast convergence rate and high solution precision. The simulation results indicate that not only the proposed method is feasible but also it is better than the indirect method, which is a most popular approach for solving the optimal endo-atmospheric ascent problem. Given the same degree of solution accuracy, the new method consumes quite less time on the CPU than that of the indirect method.

The new optimization approach has high level of both solution accuracy and efficiency. It can be used in rapid trajectory designing, on-line trajectory planning and closed-loop guidance of launch vehicles. Also, the proposed Gauss pseudo-spectral method in this paper is a new and efficient method for solving general TPBVPs.

The paper provides a new hybrid optimization method for rapid endo-atmospheric ascent trajectory planning of launch vehicles.

According to the study of the space flight market, there is a demand for space suborbital flights including commercial tourist flights. However, one of the challenges is to design a mission and a vehicle that could offer flights with relatively low G-loads. The project of the rocket-plane in a strake-wing configuration was undertaken to check if such a design could meet the FAA recommendation for this kind of flight. The project concept assumes that the rocket plane is released from a slowly flying carrier plane, then climbs above 100 kilometers above sea level and returns in a glide flight using a vortex lift generated by the strake-wing configuration. Such a mission has to include a flight transition during the release and return phases which might not be comfortable for passengers. Verification if FAA recommendation is fulfilled during these transition maneuvers was the purpose of this study.

The project was focused on the numerical investigation of a possibility to perform transition maneuvers mentioned above in a passenger-friendly way. The numerical simulations of a full-scale rocket-plane were performed using the simulation and dynamic stability analyzer (SDSA) software package. The influence of an elevator deflection change on flight parameters was investigated in two cases: a transition from the steep descent at high angles of attack to the level glide just after rocket-plane release from the carrier and an analogous transition after re-entry to the atmosphere. In particular, G-loads and G-rates were analyzed.

As a result, it was found that the values of these parameters satisfied the specific requirements during the separation and transition from a steep descent to gliding. They would be acceptable for an average passenger.

To verify the modeling approach, a flight test campaign was performed. During the experiment, a rocket-plane scaled model was released from the RC model helicopter. The rocket-plane model was geometrically similar only. Froude scales were not applied because they would cause excessive technical complications. Therefore, a separate simulation of the experiment with the application of the scaled model was performed in the SDSA software package. Results of this simulation appeared to be comparable to flight test results so it can be concluded that results for the full-scale rocket-plane simulation are also realistic.

It was proven that the rocket-plane in a strake-wing configuration could meet the FAA recommendation concerning G-loads and G rates during suborbital flight. Moreover, it was proven that the SDSA software package could be applied successfully to simulate flight characteristics of airplanes flying at angles of attack not only lower than stall angles but also greater than stall angles.

The application of rocket-planes in a strake-wing configuration could make suborbital tourist flights more popular, thus facilitating the development of manned space flights and contributing to their cost reduction. That is why it was so important to prove that they could meet the FAA recommendation for this kind of service.

The original design of the rocket plane was analyzed. It is equipped with an optimized strake wing and is controlled with oblique, all moving, wingtip plates. Its post-stall flight characteristics were simulated with the application of the SDSA software package which was previously validated only for angles of attack smaller than stall angle. Therefore, experimental validation was necessary. However, because of excessive technical problems caused by the application of Froude scales it was not possible to perform a conventional test with a dynamically scaled model. Therefore, the geometrically scaled model was built and flight tested. Then a separate simulation of the experiment with the application of this model was performed. Results of this separate simulation were compared with the results of the flight test. This comparison allowed to draw the conclusion on the applicability of the SDSA software for post-stall analyzes and, indirectly, on the applicability of the proposed rocket-plane for tourist suborbital flights. This approach to the experimental verification of numerical simulations is quite unique. Finally, a quite original method of the model launching during flight test experiment was applied.

The purpose of this paper is to design a global robust and continuous control scheme for the attitude tracking control problem of the reentry vehicle with parameter uncertainties and disturbances.

First, feedback linearization is applied to the model of reentry vehicle, resulting in three independent uncertain subsystems. Then a new second-order time-varying sliding function is proposed, based on which a continuous second-order time-varying sliding mode control (SOTVSMC) law is proposed for each subsystem. The global robustness and convergence performance of the closed-loop reentry vehicle control system under the proposed control law are proved.

Simulation is made for a reentry vehicle through the assumption that there is external disturbance to aerodynamic moment and the aerodynamic parameters as well as the atmospheric density are perturbed. The results verify the validity and robustness of the proposed strategy.

The SOTVSMC attitude controller based on feedback linearization is proposed for the reentry vehicle. The advantages of the proposed SOTVSMC are twofold. First, the global second order sliding mode is established, which implies that the closed-loop system is global robust against matched parameter uncertainties and disturbances in reentry. Second, the chattering problem is significantly alleviated.