Hypersonic Vehicle Research Articles

Evaluation of Various Flow Control Methods in Reducing Drag and Aerodynamic Heating on the Nose of Hypersonic Flying Objects

Effective deduction of air heating load and drag is a critical issue in hypersonic vehicle engineering applications. In this research, seven various geometrical models have been proposed to study and compare the effect of each configuration on the flow field, drag, and aerodynamic heating deduction under the same flow conditions. The presented configurations in this study: (a) blunt-body geometry as a reference of comparison, (b) blunt-body geometry with a spike, (c) blunt-body geometry with an counter flow jet, (d) blunt-body geometry with a spike and counter flow jet, (e) blunt-body geometry with a spike and aerodisk, (f) blunt-body geometry with a spike, aerodisk, and root counter flow jet, (g) blunt-body geometry with a spike, four aerodisks and root counter flow jet. The Reynolds-Averaged equations have been solved using the Finite Volume Method (FVM) along with the shear stress turbulence model (k-ω SST). The flow is assumed compressible, steady-state, and axisymmetric with a free stream Mach number of 6. According to the study of each configuration’s performance related to the parameters of drag, maximum pressure, and maximum heat flux factors on the blunt-body walls, (g) configuration with a drag factor of 0.2699, maximum pressure factor of 209.8, and maximum heat flux factor of 25.1, has the most deduction on the blunt-body walls among the seven configurations. The deduction percentage of drag, maximum pressure, and maximum heat flux factors of (g) configuration to (a) configuration are %72.1, %94.5, and %79.9, respectively, which significantly diminished drag and heat flux. Also, the best configuration scenarios for drag and aerodynamic heating deduction are geometrical models of g, f, d, e, c, b, and a, respectively.

Research on the cooling performance of the discontinuous transpiration surface structure for the leading edge of a hypersonic vehicle

Transpiration cooling is an efficient thermal protection technology that can be applied to hypersonic aircraft. This paper takes the leading edge transpiration cooling system of aircraft as the research object, the full-field coupling numerical method is mainly used to study the cooling performance of the transpiration cooling system of the aircraft leading edge with a flight altitude of 20 km and a flight Mach number of 6. The correctness and accuracy of the numerical method and results are validated by comparison with experimental data in reference. The effects of 5 different types of coolant gases and injection rates on the cooling effectiveness of a transpiration cooling system with a single transpiration surface arrangement are first studied. The results show that helium gas with a larger specific heat capacity has better cooling performance. As the coolant injection rate increases, the average temperature of the solid wall decreases, but the coolant will be wasted. In order to improve the utilization rate of coolant and achieve system weight reduction and large area thermal protection, solve the problem of uneven surface temperature and high stagnation point temperature of the porous media, a discontinuous transpiration surface structure is designed, and the effectiveness of varying distances between the two transpiration surfaces and the combination of varying porosity on the cooling performance of the discontinuous transpiration surface structure are studied. The results show that an increase in the distance between transpiration surfaces can reduce the coolant flow rate at the stagnation point, leading to a slight increase in the temperature at the stagnation point and a decrease in cooling efficiency of about 10 %. At the same time, the temperature at the rear wall of the second transpiration surface decreases and the average efficiency is improved. When the variable porosity of the discontinuous transpiration surface is applied, the stagnation point temperature is reduced from 1218.17 K to 486.72 K, the cooling efficiency at the stagnation point is greatly improved, and the maximum reaches 0.874.

Three-dimensional numerical study on reducing hypersonic blunt body drag and aeroheating with spike-aerodisk-bleed air channel

Shock drag and aeroheating are major problems faced by hypersonic vehicles. This study proposes a novel three-dimensional model combining spike-aerodisk with a bleed air channel featuring multiple lateral nozzles for drag and aeroheating reduction. The high-temperature and high-pressure air at the stagnation point of the aerodisk is introduced into the channel and ejected laterally through multiple nozzles to modify the flow field structure and reduce drag and aeroheating. In contrast to previous research employing a two-dimensional axisymmetric assumption, our innovation considers the structure of the exit nozzles and the spacing between nozzles. Due to multiple lateral nozzles, the connection thickness between nozzles can lead to asymmetric flow phenomena. Three-dimensional flow field simulation has been employed with the Reynolds-averaged Navier-Stokes (RANS) equations coupled with the shear stress transport (SST) k-ω turbulence model. The effects of the convergence half-angle of the channel inlet, the divergent angle of the nozzle, and the nozzle position on modifying the characteristics of the flow field are studied. Increasing the convergence half-angle of the channel inlet can further push away the reattachment shock, but when it exceeds 60°, the drag reduction performance degrades. Reducing the divergent angle of the nozzle and increasing the number of nozzles can reduce the drag coefficient and the peak value of the Stanton number. However, when the divergent angle is reduced to 10°, the performance of reducing drag and aeroheating is weakened. In the range of research parameters, when the convergence half-angle is 60°, the divergent angle is 15°, and the nozzles are located in the middle of the spike, compared with the plain spike-aerodisk model, the drag coefficient and peak Stanton number of the combined model are further reduced by 17.8 % and 34.8 %, respectively. Our findings confirm that a multi-nozzle can effectively reduce drag and aeroheating while the channel structure is viable in practice.

Thermal conductivity prediction of aerogel considering heat treatment effect and the application to a TPS of long endurance HFVs

Purpose For a thermal protection system (TPS) of long endurance hypersonic flight vehicle (HFV), its thermal insulation property not only determines by the manufactured morphology but also changes along time. A thermal conductivity prediction model for aerogel considering heat treatment effect is carried out and applied to solve the heat conduction problem of a TPS. The aim of this study is to provide theoretical and numerical references for further development of aerogels applying to TPSs. Design/methodology/approach A thermal conductivity prediction model for aerogel is established considering treatment effect. The heat conduction problem of a TPS is derived and solved by combining the differential quadrature method and the Runge–Kutta method. The prediction results of aerogel thermal conductivities are verified by comparing with those in literature, while the calculated temperature field of TPS is verified by comparing with that by ABAQUS. Findings Numerical results show that when applying the current prediction model, the calculated high temperature area in the aerogel layer is narrowed due to the decrease of the thermal conductivity during heat treatment process. Originality/value This study will be beneficial to carry out the precise design of TPS for long endurance HFVs.

Unmeasurable flexible dynamics monitoring and tracking controller design for guidance and control system of hypersonic flight vehicle

The guidance and control system (GCS) is an important and indispensable component of hypersonic flight vehicle (HFV), and it can be viewed as a special cyber–physical system (CPS). However, the monitoring of the HFV dynamics is difficult since the inertial navigation equipment can only provides the information of acceleration and angular rate. In this case, the flexible states of HFV are difficult to be measured. This paper proposes a sliding mode observer (SMO) based GCS design method for HFV. With considering the measurement difficulties, the flexible states of HFV are assumed to be unmeasurable. Then a SMO is designed to estimated them. Combined with the actual flight environment of HFV, the measuring devices are assumed to be suffering from disturbance, then the SMO should not only guarantee the estimation error stable, but also reduce the influence of external disturbances. What is worse, the input of HFV is suffering from input delay, which brings more difficulties to the controller design. By state transformation,input delay model of HFV is transformed into a delay free model,then a SMO based robust controller is proposed for HFV based on the stable inversion controller design strategy. By the proposed SMO based controller, the GCS can be constructed for HFV. Numeral simulation are carried on HFV, and simulation results show shat, the designed SMO can realize the monitoring of flexible states of HFV and the designed GCS can get a good tracking performance.

Numerical study on the non-equilibrium characteristics of high-speed atmospheric re-entry flow and radiation of aircraft based on fully coupled model

The strong coupling interactions of non-equilibrium flow, microscopic particle collisions and radiative transitions within the shock layer of hypersonic atmospheric re-entry vehicles makes accurate prediction of the aerothermodynamics challenging. Therefore, in this study a self-consistent non-equilibrium flow, collisional–radiative reactions and radiative transfer fully coupled model are established to study the non-equilibrium characteristics of the flow field and radiation of vehicle atmospheric re-entry. The comparison of the present calculation results with flight data of FIRE II and previous results in the literature shows a reasonable agreement. The thermal, chemical and excited energy level non-equilibrium phenomena are obtained and analysed for the different FIRE II trajectory points, which form the critical basis for studying the heat transfer and radiation. The non-equilibrium distribution of excited energy levels significantly exists in the post-shock and near-wall regions due to the rapid vibrational dissociation and electronic under-excitation, as well as the wall catalytic reactions. The analysis of stagnation-point heating of FIRE II illustrates that the translational–rotational convection and the dissociation component diffusion play key roles in the aerodynamic heating of the wall region. The spectrally resolved radiative intensity in the entire flow field indicates that the vacuum ultraviolet radiation caused by the high-energy nitrogen atomic spectral lines makes the main contribution to the radiative transfer. Finally, it is found that the non-equilibrium flow–radiation coupling effect can exacerbate the excited energy level non-equilibrium, and further affect the gas radiative properties and radiative transfer. This fully coupled study provides an effective method for reasonable prediction of atmospheric re-entry flow and radiation fields.

Conformal Fabrication of Thick Film Platinum Strain Gauge Via Error Regulation Strategies for In Situ High-Temperature Strain Detection.

Monitoring high-temperature strain on curved components in harsh environments is a challenge for a wide range of applications, including in aircraft engines, gas turbines, and hypersonic vehicles. Although there are significant improvements in the preparation of high-temperature piezoresistive film on planar surfaces using 3D printing methods, there are still difficulties with poor surface compatibility and high-temperature strain testing on curved surfaces. Herein, a conformal direct ink writing (CDIW) system coupled with an error feedback regulation strategy was used to fabricate high-precision, thick films on curved surfaces. This strategy enabled the maximum amount of error in the distance between the needle and the substrate on a curved surface to be regulated from 155 to 4 μm. A conformal Pt thick-film strain gauge (CPTFSG) with a room-temperature strain coefficient of 1.7 was created on a curved metallic substrate for the first time. The resistance drift rate at 800 °C for 1 h was 1.1%, which demonstrated the excellent stability and oxidation resistance of the CPTFSG. High-temperature dynamic strain tests up to 769 °C revealed that the sensor had excellent high-temperature strain test performance. Furthermore, the CPTFSG was conformally deposited on an aero-engine turbine blade to perform in situ tensile and compressive strain testing at room temperature. High-temperature strain tests were conducted at 100 and 200 °C for 600 and 580 με, respectively, demonstrating a high steady-state response consistent with the commercial high-temperature strain transducer. In addition, steady-state strain tests at high temperatures up to 496 °C were tested. The CDIW error modulation strategy provides a highly promising approach for the high-precision fabrication of Pt thick films on complex surfaces and driving in situ sensing of high-temperature parameters on curved components toward practical applications.

Phase transition cooling structure towards the wing antenna of hypersonic vehicle

In order to optimize the heat dissipation efficiency of the wing antenna of a hypersonic vehicle, a comprehensive performance-based investigation of active cooling of the wing antenna unit is proposed, taking into account multiple factors. The volume of fluid (VOF) model is employed to simulate and examine the behavior of bubbles and thermal performance on the model surface with respect to the pipe inlet, pipe outlet, and pipe spacing. Furthermore, the heat dissipation efficiency of the wing antenna structure is validated using a quartz lamp as the heat source. The comprehensive performance factor is able to balance the relationship between pressure drop and heat transfer. The simulated heat dissipation efficiency shows a high level of agreement with experimental testing. These results indicate that an appropriate coolant flow rate and flux can effectively provide the degree of phase change, and vapor can only provide effective heat dissipation within a suitable range. Conversely, in situations with excessive or insufficient bubbles, heat cannot be effectively dissipated, thereby reducing the heat dissipation efficiency.

Direct Numerical Simulation of High-Enthalpy Turbulent Boundary-Layer Flow with Light Gas Injections

Two light gases (He and H2) are, respectively, introduced upstream of a high-enthalpy turbulent flat-plate flow with a boundary-layer edge Mach number of Maδ=2.25 and temperature of Tδ=1800 K. The flow condition refers to the after-shock wave flow on a blunt-body hypersonic vehicle (Duan and Martín, AIAA Journal, Vol. 49, No. 1, 2011, pp. 172–184). Direct numerical simulation results show that the injection of these light gases has little effect on the mean velocity profiles but significantly reduces the near-wall density and skin friction. The separation and fragmentation of near-wall vortical structures are restrained, and the Reynolds shear stress decreases. The injection of inert He inhibits the dissociation reaction of O2 and weakens the chemical nonequilibrium effect, resulting in enhanced mean and fluctuating temperatures. The injection of active H2 promotes the reaction between H2 and O2, which increases the mean temperature but inhibits its fluctuation. After decomposing the mean skin friction into physics-informed contributions, both injections largely reduce the turbulent kinetic energy production term Cf,T and the spatial growth term of the flow, Cf,G, through lowering the near-wall density and reducing vortices.

Distributed Formation Control with Obstacle and Collision Avoidance for Hypersonic Gliding Vehicles Subject to Multiple Constraints

Multiple hypersonic gliding vehicles’ (HGVs’) formation control problems with obstacle and collision avoidance are investigated in this paper, which is addressed in the stage of entry gliding. The originality of this paper stems from the formation control algorithm where constraints of dynamic pressure, heating rate, total aerodynamic load, control inputs, collision avoidance, obstacle avoidance, and the terminal states are considered simultaneously. The algorithm implements a control framework designed to be of two terms: distributed virtual controller and actual control input solver. The distributed virtual controller is based on distributed model predictive control with synchronous update strategy, where the virtual control signals are derived by the optimization simultaneously at each time step for each HGV under directed communication topology. Subsequently, according to the virtual control signals obtained, a coupled nonlinear equation set is solved to get actual control signals: each HGV’s bank angle together with the angle of attack. The actual control input solver adopts a feasible solution process to calculate the actual control signals while dealing with constraints. Finally, extensive numerical simulations are implemented to unveil the proposed algorithm’s performance and superiority.

Infrared radiation characteristics of dagger-type hypersonic missile

Hypersonic vehicles emit strong infrared radiation from their high-temperature exhaust plume and body, which is critical for infrared early warning, tracking, and guidance. In this work, a comprehensive analysis is conducted on the factors involved in air dissociation reaction within the shock layer of hypersonic missile heads, as well as the multi-component afterburning effect of the exhaust plume. A novel Reverse Monte Carlo Method (RMCM) is proposed for infrared radiation calculation, which utilizes two-dimensional Low-Discrepancy Sequences (LDS) to improve computational accuracy. The numerical calculations for a dagger-type missile show that afterburning reactions increase the temperature on the centerline of the outlet exhaust plume by about 1000 K. The total infrared radiation intensity of the missile is the highest in the 1–3 μm band, with the high-temperature wall of the nozzle being the primary source of solid radiation, and gas radiation primarily coming from H2O. The radiation intensity of the missile exhaust plume in the 3–5 μm band is the highest, with radiation sources primarily coming from CO2, CO, and HCl. Afterburning reactions of the exhaust plume increase the total infrared radiation intensity of the missile by about 0.7 times. These results can provide reference for the detection and guidance of hypersonic missiles.

Modified deep deterministic policy gradient based on active disturbance rejection control for hypersonic vehicles

Modified deep deterministic policy gradient based on active disturbance rejection control for hypersonic vehicles

Multiscale Interpenetrated/Interconnected Network Design Confers All-Carbon Aerogels with Unprecedented Thermomechanical Properties for Thermal Insulation under Extreme Environments.

With ultralight weight, low thermal conductivity, and extraordinary high-temperature resistance, carbon aerogels hold tremendous potential against severe thermal threats encountered by hypersonic vehicles during the in-orbit operation and re-entry process. However, current 3D aerogels are plagued by irreconcilable contradictions between adiabatic and mechanical performance due to monotonicity of the building blocks or uncontrollable assembly behavior. Herein, a spatially confined assembly strategy of multiscale low-dimensional nanocarbons is reported to decouple the stress and heat transfer. The nanofiber framework, a basis for transferring the loading strain, is covered by a continuous thin-film-like layer formed by the aggregation of nanoparticles, which in combination serve as the fundamental structural units for generating an elastic behavior while yielding compartments in aerogels to suppress the gaseous fluid thermal diffusion within distinct partitions. The resulting all-carbon aerogels with a hierarchical cellular structure and quasi-closed cell walls achieve the best thermomechanical and insulation trade-off, exhibiting flyweight density (24mg cm-3 ), temperature-constant compressibility (-196-1600 °C), and a low thermal conductivity of 0.04829W m-1 K-1 at 300 °C. This strategy provides a remarkable thermal protection material in hostile environments for future aerospace exploration.

Hypersonic vehicle trajectory prediction based on temporal convolutional network

Hypersonic vehicle trajectory prediction based on temporal convolutional network

Hypersonic vehicle terminal velocity improvement considering ramjet safety boundary constraint

The air-breathing ramjet is an essential power plant to improve the terminal velocity of hypersonic vehicles. The high-performance requirement of ramjet makes it operate close to the safety boundary, and the optimization of the terminal velocity of vehicle is faced with challenges. An integrated flight/ propulsion model considering thrust and safety performance is constructed, and the trajectory optimization method for the whole flight phase operating process of hypersonic vehicles with multiple performance optimization indicators and complex safety constraints is studied. Moreover, the trajectory optimization results of the whole flight phase for CAV-H and X-43A (powered/unpowered) plans are compared and analyzed, and the dive-acceleration-climb strategy can improve the maneuverability of the vehicle in the ramjet operating phase. Furthermore, the terminal flight velocity of the vehicle in the powered plan is increased from 363.60 m/s to 943.57 m/s under the effect of ramjet. In addition, the trajectory optimization results of vehicle with different geometrical configurations of the ramjet are compared under safety constraints. The results show that the variable geometry ramjet enables the vehicle to increase its terminal velocity by up to 9.73 % with respect to the fixed geometry engine. Hypersonic vehicle terminal velocity is effectively improved by considering ramjet variable geometry adjustment.

Rapid ascent trajectory optimization of rocket-based combined cycle hypersonic vehicle

The ascent trajectory design of a rocket-based combined cycle (RBCC) hypersonic vehicle has many typical characteristics, including the complex working mode of the RBCC power system, strong coupling between thrust and flight state, strong nonlinear model and many complex constraints etc. This paper proposes a rapid trajectory optimization method based on sequential convex optimization to optimize the complex ascent trajectory with the RBCC power system. Firstly, a mathematical model for optimizing the ascent trajectory of the RBCC hypersonic vehicle is established to describe the angle of attack control system with and without the second-order lag. Then, the optimization model is made convex and discretized based on the convex optimization theory, and a trajectory optimization strategy is improved. Finally, taking the maximum terminal mechanical energy as optimization objective, the ascent trajectory optimization is simulated in cases with and without the second-order lag in the angle of attack control system. The simulation results show that the proposed mathematical model and the ascent trajectory optimization method are rapid and reliable and that the optimization results meet the characteristics of the RBCC power system and provide reference for the application of a RBCC power system and the design of an angle of attack control system.

Shock interactions and heating predictions on a V-shaped blunt leading edge at Mach 6–12

V-shaped blunt leading edges (VSBLEs) are usually found at the inlet lips of air-breathing hypersonic vehicles and irregular shape flows. In this work, the VSBLE flows are investigated using numerical simulations and theoretical analysis from Mach 6 to Mach 12. The simulation results show that complex shock–shock interactions around the VSBLE cause extremely high heat flux peaks, which nonlinearly increase with the freestream Mach numbers. To theoretically study the flow mechanism, the shock interactions are divided into large-scale primary shock interactions (PSIs) and micro-scale secondary shock interactions (SSIs). The PSIs are constant, but the SSIs experience a transition from Mach reflection to regular reflection with the Mach number increasing. A transition criterion for the SSIs is established by the shock interaction theory. Furthermore, the increase in the heat flux peaks is proved to be caused by the SSI transition. A semi-empirical heat flux prediction method that relates the shock intensity and heat flux amplification is established. Finally, the transition criterion and the heat flux prediction method are verified by simulations at higher Mach numbers and experiments of VSBLEs with different geometric parameters. This paper develops a theoretical analysis approach for quickly predicting the shock interaction types and heat flux peaks of the VSBLEs under a wide range of Mach numbers.

Mechanism of width effects on cavity-induced transitions at hypersonic speed using direct numerical simulation

Cavities on the surfaces of hypersonic vehicles cannot be avoided easily. Moreover, they can trigger boundary layer transition under certain conditions. However, little progress has been reported on boundary layer transitions induced by a three-dimensional (3-D) shallow cavity. In this study, transitions induced by six 3-D shallow cavities with the same length–depth ratios of 20 and different widths of 0.25, 0.5, 0.75, 1.0, and 1.5 times the baseline width (W), as well as infinite width, were investigated. A direct numerical simulation was conducted using the Roe scheme with 4th-order minimum dispersion and controllable dissipation, and weighted essentially non-oscillatory reconstruction, based on our in-house code: Unsteady NavIer–STokes equations solver. Cavity width was observed to have non-monotonic influences on transition. Both the 0.25 and 1.0 W cavities could induce transition constantly. Moreover, flow was maintained as laminar past the 1.5 W and InfW cavities. For the 0.5 and 0.75 W cavities, intermittent transition was observed with different intermittency factors. The intermittent transition phenomenon was determined to be caused by the periodic increase and decrease in the adverse pressure gradient (APG) in the front part of cavity. Notably, the recirculation with a synchronic size change was the origin of the APG oscillation.

An Intelligent Autonomous Morphing Decision Approach for Hypersonic Boost-Glide Vehicles Based on DNNs

In addressing the morphing problem in vehicle flight, some scholars have primarily employed reinforcement learning methods to make morphing decisions based on task. However, they have not considered the constraints associated with the task process. The innovation of this article is that it proposes an intelligent morphing decision method based on deep neural networks (DNNs) for the autonomous morphing decision problem of hypersonic boost-glide morphing vehicles under process constraints. Firstly, we established a dynamic model of a hypersonic boost-glide morphing vehicle with a continuously variable sweep angle. Then, in order to address the decision optimality problem considering errors and the heat flux density constraint problem during the gliding process, interference was introduced to the datum trajectory in segments. Subsequently, re-optimization was performed to generate a trajectory sample library, which was used to train an intelligent decision-maker using a DNN. The simulation results demonstrated that, compared with the conventional programmatic morphing approach, the intelligent morphing decision maker could dynamically determine the sweep angle based on the current flight state, leading to improved range while still adhering to the heat flux density constraint. This validates the effectiveness and robustness of the proposed intelligent decision-maker.

Progress on prediction of multiscale coupling thermal effects of hypersonic vehicles

Progress on prediction of multiscale coupling thermal effects of hypersonic vehicles