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An experimental investigation was carried out to find the effect of a strut spanning across the width of a single expansion ramp of a laboratory model scramjet thruster as a thrust vector control device. Cold flow tests were conducted with operating total pressures ranging from 4 bar to 10 bar and the thruster exhaust was to ambient atmosphere. The mass flow rate varied from 0.105 kg/s to 0.263 kg/s. Experiments were conducted by varying the strut height at different operational total pressures to find if any thrust vector control could be achieved to supplement the maneuverability of hypersonic vehicle with aerodynamic control. The laboratory model consisted of an isolator, a divergent combustor followed by a single expansion ramp. Except the side walls, the thruster was fabricated with stainless steel. A high quality acrylic sheet was used for internal flow visualization by a schilieren system. The wall pressure was recorded at different locations from the combustor inlet to ramp trailing edge. Shock pattern was studied from the schilieren images and it was observed that an increase in strut height caused a downward deflection of the exhaust. From the wall pressure distribution, two dimensional side force coefficient and pitching moment coefficient were calculated and the effect of strut height variation on the above coefficients was plotted. Results from experiments indicated that the presence of the strut yielded noticeable changes in side force and pitching moment. The increase in strut height provided exhaust stream directional changes which may be useful in maneuvering the vehicles employing scramjet propulsion system.

This paper deals with the anti-windup control problem for a class of feedback linearizable systems subject to external disturbances and constrained actuators. The original system is approximately feedback linearized into a tractable form with additive mismatched disturbances. A baseline disturbance suppression control is built on the combination of a disturbance observer and a compensation law. Moreover, this baseline control is strengthened by two anti-windup mechanisms to accommodate the possible input saturation, leading to the anti-windup disturbance suppression control. The resulting control schemes guarantee output tracking errors to be bounded in the presence of input saturation, while the zero steady-state tracking is finally achieved due to the recovery property. The proposed methods are applied to the longitudinal tracking control of hypersonic flight vehicles to highlight their efficiency and superiority.

This paper proposes a type-2 fuzzy adaptive sliding mode control scheme for tracking control of hypersonic aircraft with uncertainties. This method uses full-state feedback to linearize the nonlinear model of hypersonic aircraft. Combining the interval type-2 fuzzy approach and adaptive sliding mode control keeps the system stable in the existence of uncertain parameters. For rapid stabilization of the system, the adaptive laws are designed using a direct constructive Lyapunov analysis together with a well-established type-2 fuzzy logic control. Simulation test results indicate that the proposed control scheme provides enhancement of robustness to parametric uncertainty and improvement in tracking performance of the hypersonic aircraft.

This study concerns with the attitude and velocity tracking control problem for the longitudinal model of hypersonic flight vehicles, which is nonlinear in aerodynamics with model uncertainties and external disturbances. By employing back stepping sliding mode method and the strictly-lower-convex-function-constructing nonlinear disturbance observer (SNDOB), a novel composite controller is proposed to guarantee the system tracking error to converge to a small region containing the origin. Besides, several proper adaptive laws are also introduced to make the controller avoid of the differential explosion problem and be chatter-free. Compared with other robust flight control approaches, key novelties of the developed method are that one new SNDOB is proposed and drawn into the virtual control laws at each step to compensate the disturbances and that adaptive laws are utilized to simplify the tedious and complicated differential operations. Finally, it is demonstrated by the simulation results that the new method exhibits not only an excellent robustness but also a better disturbance rejection performance than the convention approach.

At hypersonic speed, strong nonlinear couplings will bring enormous challenges to the attitude control of a hypersonic flight vehicle. To reduce the negative impacts of the couplings, a hierarchical coordinated controller is proposed for a generic hypersonic vehicle in this paper. By using a statistical sampling method, a series of coupling analyses between any two different attitude variables groups are studied firstly, and the matrices that describe the coupling degrees are provided. Based on the coupling degree matrices, two coordinated controllers are designed for the attitude angles and angular rates, respectively. The simulation results indicate that the proposed controller can effectively coordinate the coupling impacts, and achieve smooth attitude tracking.

In this paper, the related research and analysis are made on the finite-time tracking control of hypersonic aircrafts. Two finite-time anti-saturation integral sliding mode controllers are designed with the consideration of external disturbances, coupling responses, model parameter uncertainties, and input saturation. Firstly, based on the mean value theorem, the control problem with input saturation constraints is transformed into the one without input constraints, which overcomes the assumption of the input boundedness. Secondly, two adaptive finite-time integral sliding mode controllers are designed: the first controller is finite-time stable, while the second controller is practical finite-time stable which utilizes the adaptive algorithms with low-pass filtering properties to ensure the boundedness of adaptive parameters. Finally, the digital simulation is carried out on different types of signals to further verify the validity and robustness of the proposed finite-time anti-saturation controllers.

Methods are reported for less computationally expensive and more accurate implementations of the direct simulation Monte Carlo (DSMC) method for the simulation of high speed gas flows over arbitrarily shaped bodies. A new particle-tracking algorithm with a saving of computational time of up to 10% is reported in which tracking of particles is done with the help of big triangles having vertices lying on the boundary curves. An algorithm has been developed to generate DSMC cells for collision and sampling that contain a prescribed number of molecules. This algorithm is able to generate over 90% cells having the optimum number of seven or eight molecules for simulating collisions. Sampling for macroscopic properties is done on dynamic cells that contain a number of particles varying spatially as a function of the local number density. A criterion for finding the number of particles in sampling cells is presented. This criterion has been found to result in accurate and fast simulation of two-dimensional hypersonic flows of argon over a wedge, and argon and nitrogen over a circular cylinder.

A ballistic error propagation algorithm for glide trajectories of a hypersonic glide vehicle is originally proposed based on the perturbation theory. Perturbation equations that treat perturbations as external inputs and flight state deviations as state variables are established. Based on the reasonable simplification assumptions, the analytic expression of the state transition matrix is deduced and thus the ballistic error propagation model is established. A transposed coordinate frame is introduced to simplify the development of the perturbation equations and the error propagation model. By employing the gravity anomaly as the single perturbation factor, the proposed algorithm is validated and verified by numerical experiments. When compared with the traditional method, the proposed method takes only just a quarter computational costs and restrains the method errors beneath 10% of the total terminal deviations. It is an effort that initially focuses on the error propagation mechanism of the glide trajectory and the proposed model has sufficient precision for the analysis of modeling deviations, thus laying the foundation of correcting the modeling deviations and enhancing the accuracy of vehicle’s flight states.

By using light-weighted material in hypersonic flight vehicle, the vehicle body can be easily deformed. The mutual couplings in aerodynamics, flexible structure and control system will result in aeroservoelasticity problem. Based on the six-DOF geometry of a hypersonic flight vehicle, this paper uses assumed modes method, Shock-Expansion theory and Piston theory to develop the structural elastic model and the aerodynamic model, with actuator incorporated, the open-loop aeroservoelastic model of a six-DOF HFV is established. In open-loop analysis, the response of the elastic coordinates to the tail rudder deflection and the frequency characteristic of the transfer function of the tail rudder deflection to the first elastic coordinate are studied. A high gain feedback controller is designed for the aeroservoelastic model, and closed-loop analysis is conducted based on this model. The simulations show the influence of the elastic motion on the aerodynamic forces and moments.

An adaptive decoupling control approach using fuzzy-neural network (FNN) observer for a class of MIMO nonlinear systems with parameter uncertainties is presented in this article. First, a decoupling controller is constructed based on the decentralized control theory. Furthermore, the system coupling terms and uncertainties are estimated by the FNN observer and added into the control law for compensation. The FNN approximate-matrix update law and the control law guarantee that the tracking errors of the system states, the observer states and the approximate matrix are all uniformly ultimately bounded within a region that can be kept arbitrarily small. Secondly, a model for the hypersonic vehicle is given and an attitude controller is designed using the decoupling control approach. Finally, simulations are carried out on the hypersonic vehicle to demonstrate the effectiveness of the proposed method.

The presence of system strong constraints and large uncertainties in reentry hypersonic vehicles deteriorate the system performance. This paper considers disturbance observer-based sliding manifold predictive control design for a reentry hypersonic vehicle. In the outer loop, a new sliding manifold-based predictive control is proposed to realize reference command tracking. In the inner loop, a nonlinear disturbance observer is adopted to estimate lumped disturbances and a sliding manifold-based predictive control is designed to track the expected signal obtained from the outer loop. The use of sliding manifolds and disturbance observer enhance the control system’s robustness to uncertainties. The mode predictive control approach guarantees the satisfaction of the system constraints. The effectiveness of the proposed algorithm is verified by theoretical analysis and simulations.

A novel highly constrained optimal gliding guidance algorithm, which can satisfy multiple path and terminal constraints in the presence of no-fly zones for hypersonic vehicle, is investigated in this paper. It employs line-of-sight angle to describe the relative location relationship between flight path and no-fly zones, and proposes boundaries selection algorithm for each no-fly zone, which is called avoidance strategy, with minimum energy consumption. The boundaries determined by avoidance strategy divide the flight process and guidance task into several stages. In each stage, it constructs guidance models based on quasi-equilibrium glide condition in longitudinal and lateral directions respectively, and employs optimal control to generate guidance law to satisfy terminal location, altitude, and flight path angle constraints. In addition, all the path constraints, including heating rate, dynamic pressure, and overload, are converted into the angle-of-attack constraint to ensure the guidance mission can be accomplished successfully. This algorithm is independent of standard trajectory, and all the guidance commands, including the bank angle and the angle of attack, can be calculated analytically in real time and adaptively. Finally, the simulation results of CAV-H indicate that the proposed strategy can guide the vehicle to satisfy multiple different constraints with high precision and avoid no-fly zones effectively.

A nonlinear robust control strategy is proposed for flexible hypersonic vehicles with aerodynamic parameters uncertainties. Firstly, the longitude dynamics of this vehicle are divided into velocity subsystem, altitude and flight path angle subsystem, angle of attack, and pitch rate subsystem. Then the smooth second order sliding mode controller and disturbance observer are designed to ensure the finite time stability for each subsystem. Furthermore, the stability of the whole system is guaranteed via the assumption of multiple-time scale characteristic. Moreover, the effect of flexible modes on rigid-body states can be estimated exactly using the disturbance observer. In addition, two representative simulations are carried out to analyze the effect of elevator deflection and canard on the dynamics of velocity, flight path angle, and angle of attack. Finally, the Monte Carlo simulations are provided to verify the effectiveness of the proposed control strategy.

This paper presents a tracking control problem of flexible air-breathing hypersonic vehicle with input constraint and aerodynamic uncertainty. Without ignoring aero-propulsive and elevator-to-lift couplings, a control-oriented model including aerodynamic uncertainty is firstly established. Then a robust adaptive backstepping control scheme is designed, in which the control-oriented model does not need to be transformed into linear parameterization formulation. Upper bounds of the uncertain terms do not need to be known in advance, which are estimated online by designing robust adaptive laws. To further consider input constraint, a constrained robust adaptive backstepping controller is proposed to simultaneously handle input constraint and aerodynamic uncertainty. Finally, the compared simulation results show the effectiveness of the designed control strategy.

Buzz is an important issue for a scramjet engine. A mathematical model of buzz oscillations is necessary for control system design. Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model that is accurate to some extent but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model of buzz oscillations for a scramjet engine is built by introducing the modeling idea of Moore–Greitzed model for compressors. The introduction of characteristic lines avoids the complex interactions in hypersonic inlet, such as shock–shock interactions and shock–boundary layer interaction. And the inlet characteristics are obtained from the pressure signal of combustor. Based on the established buzz model, we can predict the inlet performance, characterize the stability margin of inlet, reflect the oscillatory characteristics of inlet buzz including the dominant amplitude and frequency and describe the transition process of inlet buzz.

Powerful actuators are both indispensable and critical components for the flight control system of hypersonic vehicles. More importantly, the performance of actuators has substantial impact on the control ability; therefore, when designing the control system, one needs to fully take into account actuator restraints in order to meet the efficient and precise control demands under complex flight conditions. In this article, the advanced flight control methods concerned with actuator limitations are discussed for hypersonic vehicles. First, the longitudinal model of hypersonic vehicle is established with consideration of the nonlinear coupling dynamics. Second, the actuator constraints with the effect of elastic deformation are introduced to this built model and then the resulting unstable dynamics characteristic and the control limitation conditions are analyzed for hypersonic vehicle. Furthermore, the advanced flight control laws are designed by using the differential geometry principle and the total energy theory. Finally, simulation results verify the feasibility of the proposed methods for hypersonic vehicle.

The most recent idea of using the exothermic coating to reduce the drag coefficient on blunt noses in hypersonic flows is numerically simulated to study parametrically the effect of material properties. The conjugate heat transfer is considered by simultaneous solution of flow and solid phase governing equations in numerical procedure. The results show that changing some effective parameters in material properties may increase the coating effectiveness significantly. Using such analysis leads to choose the proper materials for specific flight missions.

Dual combustion ramjet (DCR) is one of the most promising and realistic propulsion systems to realize hypersonic flight. The flow fields and the performance of the full-size DCR in Ma4/17 km and Ma6/25 km flight conditions are investigated through direct-connected experiments and numerical simulations. The pressure distributions from simulations are in agreement with that from experiments under both cold flow and hot flow conditions. Different combustion modes are revealed according to numerical results: purely subsonic flow field is established in the front part of the combustor in Mach 4 condition, and there is a thermal choked throat; only central flow is subsonic in Mach 6 condition, and the peripheral supersonic air results in a lower static temperature (less than 2000 K), which is beneficial to thermal protection of the combustor wall. The thrust increment increases with the increasing of the equivalence ratio (Φ). The thrust increment is 8.1 kN for Mach 4 when Φ = 0.9; however, further increasing the equivalence ratio causes unstart of the supersonic intake. The thrust increment is 3.15 kN for Mach 6 when Φ = 1.0. The equivalence ratio affects the combustion efficiency and the specific impulse with the same trend. The maximum combustion efficiency is 0.91 for Mach 4 and 0.89 for Mach 6. The maximum specific impulse is 13.3 kN·s/kg for Mach 4 and 7.96 kN·s/kg for Mach 6. In general, the performance is good and the DCR is worth further investigating.

The article mainly employed the numerical simulation method to conduct the 2D hypersonic inlet performance research. In this study, the numerical simulation method that includes a two-equation turbulence model was validated through the test case firstly. For the purpose of oppressing unstart phenomena, extending operating range and improving comprehensive performance of the inlet, the bleeding system and isentropic expansion arc surface were applied to the inlet, respectively. Different kinds of configurations were designed through changing the length of the bleeding or changing the radius of the arc surface. The influence of the bleeding and isentropic expansion arc surface on the 2D hypersonic inlet was analyzed systematically based on different objective parameters. The calculative results show that the bleeding system with appropriate value of the length can improve comprehensive performance of hypersonic inlet remarkably either on the design point or the off-design points, though it has a very small negative impact on the inlet in terms of the mass flow ratio coefficient. As for isentropic expansion arc surface, it can improve the inlet’s performance at the design point on the conditions that the radius of isentropic expansion arc surface has an appropriate value. But it has much more disadvantages over the basic configuration reversely when it comes to the off-design conditions or the large size of the radius of isentropic expansion arc surface.

One of the most critical challenges for turbine-based combined cycle propulsion systems utilized in hypersonic civil aircrafts (Mach 0 to Mach 5) lies in guaranteeing the propulsion system to operate efficiently and stably along a very wide operating range. This has resulted in a need to investigate the matching restraints between the inlet and turbine-based combined cycle engine and nozzle from the integration point of view at the early stage of design. To get a good understanding of this integration concept design process, in this article, several basic design principles were first put forward. Next, a fast, versatile and trusted modular analysis tool for this concept design was developed and implemented. Third, several issues on this concept were analyzed including the design issues, the mechanisms of the variable geometries as well as the control schedules. Finally, a solution for the turbine-based combined cycle propulsion system was carried out and discussed. The results show that at least 11 total variables were required to be adopted in the variable cycle propulsion system to meet the basic mission requirements for Mach 5 civil airplane. How to guarantee these variable geometries function efficiently and reliably along the wide flight regime is one of the most critical technology challenges to make this concept feasible. The uniqueness of the developed analysis tool lies in that it affords a platform capable of investigating the design constraints and the matching relationships between the various propulsion components along a wide operating range, which is beneficial to predict the potential critical technological barriers at the early design stage.