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As of now, only about 4% of apparent space has been explored. The spacecraft used to explore these areas of the unknown have to go through a design process that often does not include conditions on environmental factors. The technical project is designing a mini spacecraft, called a CubeSat, that will reconfigure into a reentry vehicle at hypersonic speeds named the Hypersonic reEntry Deployable Glider Experiment (HEDGE). The purpose of this project is to successfully collect and transmit temperature and pressure while the spacecraft is reentering the Earth’s atmosphere at hypersonic speeds. The STS research discusses the environmental effects that are caused from space exploration. The environmental effects are both directly and indirectly resulting from space mission’s successes and failures. The technical and STS research are tightly coupled as HEDGE will be a space mission that can cause environmental harm. For this stage of the technical project, the goal was to reach the critical design review phase by the end of the semester. All of the design components have been chosen to meet all the design requirements and rationale is provided to prove that the mission will succeed safely. With a successful critical design review, the project can move forward to full scale prototyping. In order to have a successful critical design review, the project got separated into six teams, each focussing on a specific subsystem. Each team was tasked with researching the best possible components and design for this mission while remaining within a mass, and cost budget. Each team also had to verify that these selected components and design would lead to a successful mission. As a result, the critical design review has been completed as a proposal suitable for NASA. This proposal describes the design choices and is a request for funding for the full scale implementation of HEDGE. Each team included their major subsystem components, rationale for choosing the parts, calculations to back up the rationale, and how their subsystem will be assembled. The STS research question examined was how continued space exploration has caused environmental issues. In order to research this question, effects of rocket emissions, reentry heating, and space debris were explored. Additionally, possible fuel alternatives were researched to reduce rocket emission harm. Furthermore, Pinch and Bjiker’s Social Construction of Technology framework was adapted to provide a potential design process to mitigate the environmental risks with space missions. In the STS thesis, one of the main topics discussed is the production of space debris from failed space missions. Space debris is constantly orbiting in close proximity with the Earth, and even pieces the size of a paint fleck can cause considerable damage to a satellite in space. This can result in a continuous rapid increase in space debris which could lead to a higher amount of space debris reentering the atmosphere which then releases harmful toxins that lead to global warming. Additionally, black carbon released from rockets has a longer effect than other vehicles’ production because rockets burn fuel higher in the atmosphere which makes the particles last much longer. Some alternatives to rocket emissions are under development, including a bacteria based fuel and a fuel created by using recycled plastics. The technical project will be ready to begin its full scale prototyping next academic year with the guidance of this year's critical design review. However, future students working on this project should focus on environmental factors surrounding this future mission, especially harm caused by the rocket used to put the CubeSat into space, and the reentry effects when the vehicle burns up in the atmosphere. Similar to the Social Construction of Technology framework used in the STS report, the future team for the technical report could follow that framework.

The design of optimal guidance law for intercepting a near-space hypersonic maneuvering target with bounded inputs is considered. Firstly, a maneuvering model for near-space hypersonic aircraft is given. Then, the aircraft acceleration prediction can be obtained using this model with two neural networks. By using the target acceleration prediction, which is taken into account when calculating the Zero Effort Miss (ZEM), an optimal sliding-mode guidance law is proposed to fulfill the guidance task. An adaptive sliding-mode switch term is designed to deal with actuator saturation and prediction errors. Finally, numerical simulations show that the proposed guidance law can reduce the energy consumption and the terminal acceleration command of the interceptor effectively.

Recent molecular-beam experiments of high-velocity atomic oxygen (O), atomic nitrogen (N), and molecular oxygen (O2) impacting carbon material at high temperature produced detailed surface chemistr...

Optimal performance is extremely important for hypersonic flight control. Different from the most existing methodologies which only consider basic control performance including stability, robustness and transient performance, this article deals with the design of nearly optimal tracking controllers for hypersonic flight vehicles (HFVs). Firstly, main controllers are developed for the velocity subsystem and the altitude subsystem of HFVs via concise fuzzy approximations. Then, optimal controllers are nearly implemented utilizing single-network adaptive critic design. Moreover, the stability of closed-loop systems and the convergence of optimal controllers are theoretically proved. Finally, compared simulation results are given to verify the superiority. The special contribution is the application of low-complex control structure owing to the critic-only network and advanced learning laws developed for fuzzy approximations, which is expected to guarantee satisfied real-time performance.

The solid-fueled Scramjet is an interesting option for supersonic combustion ramjet. It shows significant advantages such as simple fuel supply and compactness, avoiding the complex system of tanks and pipelines that encountered in liquid-fueled Scramjets. The solid-fueled Scramjet could be the simplest air-breathing engine for the hypersonic flight regime. This paper presents a comprehensive and systematic review of the research progress on solid-fueled Scramjet in various institutes and universities. It summarizes a progress overview of three types of the solid-fueled Scramjet, which covers a wealth of landmark numerical and experimental results. Based on this, several relevant key technologies are proposed. Several inherent scientific issues are refined, such as the mixing mechanism of multi-phase flow and supersonic airflow, ignition and combustion mechanism of the condensed phase in a supersonic airflow, and coupling mechanism of gas and solid phase in a supersonic flow. Finally, the historical development trend is clarified, and some recommendations are provided for future solid-fueled Scramjet.

Among flourishing control studies on air-breathing hypersonic vehicles (AHVs), actuator characteristics have not been fully accommodated. This study gives a comprehensive consideration to physical ...

SparkJet (SPJ) actuator was used to control shock and shock-on-shock interaction in high-enthalpy hypersonic flow with a total temperature of 4490 K and Mach number 6.9 for the first time. Aiming at the characteristics of rarefied gas in hypersonic flight environment, the control authority of SPJ actuator was improved by air supply method. The process of SPJ interacting with high-enthalpy hypersonic crossflow near a ramp was obtained by using high-speed schlieren camera. When pressurized cavity pressure was 133 kPa, ramp distance was 22 mm and ramp angle was 50, SPJ can eliminate 68.3 % of the whole ramp shock. Results of different experiment cases showed that, within a certain range, the control effect of SPJ on ramp shock improves with the increase of pressurized cavity pressure, ramp distance and the decrease of ramp angle. The elimination of shock by SPJ can be explained by the upward motion of the sonic line in the boundary layer due to the formation of SPJ shock, the thickening of boundary layer and the heating of local flow field. Except for the elimination of shock, the shock intersection point moves upstream and downstream corresponding to the strength changes of SPJ and SPJ shock. Accordingly, local high heat flow around the shock intersection point can be decreased. The control effect of SPJ on shock-on-shock interaction can be improved by properly increasing discharge energy.

This article presents a fuzzy adaptive tracking control algorithm for hypersonic flight vehicles with input quantization and faults. Different from the usual stabilization results, we present a method to achieve asymptotic tracking with predefined-time performance. An error transformation is established and a modified Lyapunov function (MLF) is constructed. The predefined-time asymptotic tracking control goal can be achieved by simply guaranteeing the boundness of the proposed MLF which is related to the transformed tracking errors. The proposed control strategy has three advantages. First, the predefined-time tracking performance can be ensured even when the initial tracking conditions are completely unknown. Second, no compensation is required for the fuzzy reconstruction errors to achieve accurate tracking control such that the control design is significantly simplified. Third, the derived controller is fault-tolerant, and since the system is controlled by quantized signals, the communication burden can be considerably reduced. Theoretical proof and numerical simulation are presented to demonstrate the effectiveness of the proposed control algorithm.

An adaptive dynamic programming controller based on backstepping method is designed for the optimal tracking control of hypersonic flight vehicles. The control input is divided into two parts namely stable control and optimal control. First, the back-stepping method is exploited via neural networks (NNs) to estimate unknown functions. Then, the computational load is reduced by the minimal-learning-parameter (MLP) scheme. To avoid the problem of “explosion of terms”, a first-order filter is adopted. Next, the optimal controller is designed based on the adaptive dynamic programming. In order to solve the Hamiltonian equation, NNs estimators are introduced to approximate performance indicators, achieving the approximate optimal control of hypersonic flight vehicles. Finally, the effectiveness and advantages of the control method are verified by simulation results.

In order to investigate the dynamic characteristics of blunt aircraft mounted with aerospikes and aerodisks in large-amplitude force-pitching, the Roe spatial scheme and the lower-upper symmetric Gauss-Seidel (LU-SGS) method with dual time step are employed for discretization of unsteady Navier-Stokes (N-S) equations. A parametric investigation on the flow fields is conducted by altering the pitching period, aerospike length, and aerodisk diameter consequently via a variable-controlling procedure. Dynamic characteristics of aerodynamic drag as well as the visualization of unsteady flow fields are achieved, and the results show that the aerodynamics of hypersonic aircraft under the condition of large-amplitude force-pitching vibration have hysteresis characteristics affected by periods of force-pitching vibration. In addition, when changing aerospike length and aerodisk diameter, the variation tendency of drag reduction efficiency is determined by the pitching angle of the oscillation process.

Hydrocarbon fuels offer optimal high energy per volume for scramjet applications for sustained hypersonic flight but require additional residence time due to slower ignition delays (compared with h...

Waveriders have a high lift-to-drag ratio in hypersonic states, while the flexibility of design methods must still be improved because of some deficiencies such as unsatisfactory off-design perform...

Shock tube experiments are a cornerstone of investigations into the hypersonic flight environment. However, the actual flow through the basic shock tube has not been well characterized in the literature, and nonuniformities arising in the test gases are not well understood. This work produces a quasi-one-dimensional analytical methodology for predicting the nonuniformities in the postshock test gas based exclusively on the experimentally measured history of shock speed down the tube. Wave effects extracted from shock speed variations are shown to provide all information necessary in order to reconstruct the test slug from any particular experiment. Accurate representation of flow dynamics is initially checked against argon test cases following accelerating and decelerating shock trajectories, each with a tube-end Mach number of 6.5 and a fill pressure of 66.66 Pa in a tube of 100 mm diameter. Agreement between the method and results from a viscous, axisymmetric Navier–Stokes solution is found to be within 1% in pressure and temperature.

Lateral–directional stability is a critical issue for the design of any kind of aircraft. For the hypersonic waverider, the nonaxisymmetric, flat, and slender geometric feature makes it susceptible...

Novel resin-based ablative materials are manufactured for the Thermal Protection System (TPS) of hypersonic aircraft in recent years to adapt more extreme aerothermodynamic environment. The establishment of theoretical model and simulation method for accurately predicting material's responses can guide design of TPS and significantly reduce Research and Development (R&D) cycle and cost. In order to solve the problem of inaccurate material response calculated by the heritage models, which is caused by the new mechanisms involved in new materials. On the basis of heritage physical and chemical mechanisms for resin-based ablative materials, this study further introduces the mechanisms of the local thermal non-equilibrium and thermal radiation, and a developed thermal-fluid-ablation coupled theoretical model is established. The coupled model is numerically solved by writing computer codes. After completing the verification of the model and simulation codes, the influences of the new mechanisms on the performance of ablative material are analyzed. The results show that the effect of local thermal non-equilibrium and thermal radiation cannot be ignored for current ablative materials, in which the key factors are the pore size and porosity.

To date, plasma sheath effects have not been incorporated into most CFD simulations of magnetohydrodynamics (MHD) or aerothermodynamics due to the high computational costs involved. The accurate mo...

This article analyzes the theoretical aspects of controlling the strength and thermal modes of hypersonic aircraft. The conditions for the functioning of hypersonic aircraft were also investigated, and problematic situations for their design were identified. The parameters of aerodynamic heating of surfaces and heating of thermal protection of hypersonic aircraft were estimated with an assessment of the parameters of thermal protection of hypersonic aircraft and the magnitude of the heat flux supplied to the surface, with the determination of the parameters of their thermal protection, taking into account the thermophysical characteristics of materials from thermal parameters.

Standing of an Oblique Detonation Wave (ODW) on a wedge within combustor is the prerequisite of thrust generation for ODW engine which is regarded as a novel and conceptual propulsion device with hypersonic flight Mach number. Usually a standing window of ODW is defined as the wedge angle ranged from the ODW detached angle from wedge (upper limit) to the angle that a Chapman-Jouguet (CJ) detonation occurs (lower limit). For pathological detonation cases, however, the CJ detonation cannot be achieved, and thus the lower limit of the standing window of ODW should be revisited. In present study, two types of reactions in hypersonic incoming flow that include the behavior of pathological detonation, that is, the single-step irreversible reaction with mole variation and the two-step irreversible reactions with exothermic process followed by endothermic process, have been used for studying standing behavior of ODW. The steady detonation polar analysis of ODW is carried out for both reaction systems. The results reveal that the reaction with more mole decrement and the reactions with stronger endothermic process show the pathological detonation feature and therefore modify the lower limit of standing window of ODW. Three equivalent parameters are proposed to quantitatively measure the standing window range of ODW from points of view of thermodynamics, Mach number of incoming flow and heat effect of reactions. It is found that the standing window of ODW is determined by the specific heat ratio, the overdrive degree of detonation and the endothermic level of the hypersonic incoming flow, regardless of whether the detonation is pathological or not.

Wall pressure fluctuations beneath a boundary layer are the main cause of structural vibration for hypersonic aircraft. To accurately predict the load environment of pressure fluctuations, a measur...

In this article, an integrated guidance and control design method for general nonlinear flexible hypersonic flight vehicles in the presence of dynamic uncertainties based on the [Formula: see text] adaptive state feedback control approach is presented. Initially, the 6-degree-of-freedom integrated guidance and autopilot dynamic model is organized using the combination of flexible uncertain hypersonic flight vehicle dynamic model and hypersonic flight vehicle-target relative motion model whereas aerodynamic and model uncertainties, cross-coupling effects, and disturbances are considered. The proposed integrated guidance and control method based on [Formula: see text] adaptive control scheme afterward is designed to realize control objectives including closed-loop stability, high-performance target interception, state regulation, uncertainties compensation, and reduction of the adverse flexibility effects on the overall performance. The whole Targets are proven thereupon based on the Lyapunov theory. Finally, the three-dimensional simulation results confirm the effectiveness of suggested integrated guidance and control law not only in closed-loop stability but also in high-performance target interception using the smooth control input.

In a swing-by maneuver, the variation of energy of the spacecraft is controlled by the angle, velocity, and distance of approach. These parameters are defined by the incoming trajectory, when the s...

Recent development of deep learning has shown that a deep neural network (DNN) is capable of learning the underlying nonlinear relationship between the state and the optimal actions for nonlinear o...

Hypersonic International Flight Research Experimentation 5 (HIFiRE-5) is a hypersonic flight-test experiment designed to investigate the aerothermodynamics of a 3-D geometry. The vehicle is an elli...

This paper addresses the finite time prescribed performance control problem for hypersonic flight vehicle systems with abrupt actuator faults and time-varying uncertainties. The considered actuator faults contains abrupt gain faults and saltatorial bias faults, and the unknown discontinuous variations of the fault parameters are permitted. In virtue of the finite time performance functions and error transformation, the predefined tracking performance including the convergence time, the tracking error and the overshoot for the velocity and altitude can be ensured. By adaptively estimating the lower bonds of the gain faults and the upper bounds of the bias faults, the unwanted effects of the abruptly faulty actuators can be circumvented. Meanwhile, for the purpose of compensating the unknown nonlinearities, the learning-based control strategy has been employed. Finally, a number of simulation results on the hypersonic flight vehicles are conducted to demonstrate the validity of the proposed approach.

On the basis of analytical and experimental data, this paper presents possible levels of thermal loading of reusable hypersonic aircrafts. Methods for ensuring the thermal regime of such aircrafts using various passive and active thermal protection technologies are considered. The conducted analysis of the scientific and technical background in terms of the practical implementation of these technologies allowed the authors to draw the conclusion about the relevance and expediency of creating an integrated thermal protection system for reusable high-speed aircrafts.

In this paper a nonlinear control method is proposed for the tracking control of hypersonic flight vehicles. The designed control laws do not utilize the measured flight path angle due to its inferior accuracy in practical engineering. For this, an estimated flight path angle is designed via the measurements of the altitude and velocity. A tracking differentiator is designed for constructing nonlinear disturbance observer which is used to estimate the model uncertainties including the parameter indeterminacies and external disturbances in the channels of velocity and pitch rate. A robust high-order differentiator is introduced to avoid the employment of the measured flight path angle and estimate the lumped disturbance in dynamics of flight path angle. Meanwhile, the possible saturation of the control inputs is considered and compensated by the auxiliary states. The boundness of closed-loop signals is proved through the Lyapunov theory. Comparative simulations are carried out and the results demonstrate the effectiveness of the proposed method.