Hypersonic Vehicle Research Articles

Dynamic Flow and Heat Transfer Characteristics of Uncracked Hydrocarbon Fuel under Super-Critical Pressure in the Cooling Channel of a Regeneratively Cooled Scramjet

Regeneratively cooled scramjets are successfully used as propulsion devices in hypersonic vehicles. During operation, scramjets experience acceleration. This special process causes a dynamic flow process, and heat transfer in the cooling channel commonly occurs, which may cause hazards and control difficulties for scramjets. A dynamic numerical model with a modified heat transfer coefficient calculation method was established to study the transient flow and heat transfer processes in a cooling channel. The dynamic characteristics of the flow and heat transfer under different conditions were calculated and are discussed, including the changes in the inlet fuel mass flow, heat flux, and pressure working conditions. The results indicate that the stable time of the cooling channel outlet fuel temperature is related to the rate of change in the inlet mass flow and heat flux. The stable time of the outlet fuel temperature under decreasing heat flux working conditions was approximately 12.5 s. These results summarize the dynamic flow and heat transfer characteristics, which are significant for designing cooling channels in scramjets.

Investigation of Surface-Catalycity Effects on Hypersonic Glide Vehicle Trajectory Optimization

The optimization of a hypersonic glide vehicle’s reachable domain is studied using different models to constrain the stagnation-point heat transfer. The vehicle is modeled as a high-lift common aero vehicle that uses bank-angle modulation for crossrange maneuvering with a leading-edge thermal protection system made of an ultra-high-temperature ceramic. A constrained nonlinear optimization problem is formulated to determine the range of feasible trajectories that satisfy aerodynamic heating and loading constraints. The optimal control profile is determined using a modified particle swarm optimization routine that employs a penalty function approach to handle path and terminal constraints. Baseline trajectories are determined using an existing convective heat-flux correlation, and the effect of surface catalycity on the optimized trajectories is investigated by applying a correction factor, derived from high-fidelity computational fluid dynamics simulations, to the heat-flux correlation. Results demonstrate a high sensitivity of the optimized trajectories to the underlying aerothermodynamics model such that accounting for surface catalycity expands the reachability by an order of magnitude.

Research on the Zooming Method for Determining the Flow, Heat Transfer, and Infrared Radiation of an Air-Breathing Hypersonic Vehicle Powered by a Scramjet

In order to study the infrared radiation characteristics of an air-breathing hypersonic vehicle powered by a scramjet, it is necessary to solve the internal and external flow field of the air-breathing hypersonic vehicle. Owing to the complexity and difficulty of solving the three-dimensional flow and heat-transfer process in a scramjet combustor, a quasi-one-dimensional calculation method was established. Utilizing zooming technology, a combination of quasi-one-dimensional simulation within the combustion chamber and three-dimensional numerical simulation elsewhere on the vehicle was employed to obtain the flow field. The accuracy of the zooming method in determining flow, heat transfer, and infrared radiation was verified through comparison with experimental data. The results show that under the flight condition of Ma = 6, the gas temperature and wall heat flux in the scramjet combustor first increased and then decreased along the flow direction. The Mach number of the plume was smaller than that of the free flow, while the velocity of the plume was slightly larger. In the wavelength range of 3–5 μm, as the azimuth angle increased, the integrated radiation intensity of the air-breathing hypersonic vehicle demonstrated a characteristic pear-shaped distribution.

Adaptively tracking hypersonic gliding vehicles

The motion model plays a key role in tracking hypersonic gliding vehicles (HGVs). The process noise covariance matrix is used to scale the errors of the motion model, and its design has significant influence on the tracking accuracy. In this paper, a dynamics model with an adaptive process noise covariance matrix is proposed for tracking HGVs. The dynamics model is described by a nonlinear stochastic differential equation (SDE), and its corresponding process noise covariance matrix is calculated by adopting the modified Gauss-Legendre scheme. Furthermore, an adaptive adjustment method for process noise covariances of the augmented states is designed by analyzing the variation characteristics of the model errors. Specifically, the process noise covariances are adjusted by an adaptive factor: the square of the ratio of the estimated dynamic pressure to a prior dynamic pressure. Simulation results show that the proposed model can provide higher tracking accuracy than the dynamics model with fixed process noise covariances.

Design and decoupling analysis of Thermal–Electric energy comprehensive utilization scheme based on “diamond” active cooling thermal protection system for hypersonic vehicle

High–temperature aerodynamic heat threatens the safe cruise of hypersonic vehicles, but it also contains huge energy utilization potential. In this study, the thermal–electric energy comprehensive utilization scheme was proposed to combine the energy conversion technology (ECT) with active cooling thermal protection system (ACTPS). The mass flow rate (M) varied from 0.008 kg s−1 to 0.2 kg s−1 for each channel with the inlet fluid temperature being set to 253 K. It was investigated that the leading edge of the “Enhanced” model forms longitudinal swirls. The inner and outer wall temperatures of the “Enhanced” model were lower than 275 K and 650 K, respectively, at mass flow rate of 0.076 kg s−1. Moreover, the thermoelectric performance of five different parameters (Bi2Te3 height (L1), SiGe height (L2), width for unit (H), inner wall temperature (Tin) and heat power (QT)) was explored. L1 = 1.0 mm, L2 = 0.5 mm and H=1.5 mm were selected as the optimal parameters combination for engineering applications. Within the range of QT from 0.15 W to 0.9 W, the variations range of rated current (Irc), energy conversion efficiency (ηe) and specific power (Ps) for the optimal parameters were 140 mA–460 mA, 0.69 %–2.61 % and 0.016 W g−1–0.373 W g−1, respectively.

A novel evasion guidance for hypersonic morphing vehicle via intelligent maneuver strategy

This paper presents a novel evasion guidance law for hypersonic morphing vehicles, focusing on determining the optimized wing's unfolded angle to promote maneuverability based on an intelligent algorithm. First, the pursuit-evasion problem is modeled as a Markov decision process. And the agent's action consists of maneuver overload and the unfolded angle of wings, which is different from the conventional evasion guidance designed for fixed-shape vehicles. The reward function is formulated to ensure that the miss distances satisfy the prescribed bounds while minimizing energy consumption. Then, to maximize the expected cumulative reward, a residual learning method is proposed based on proximal policy optimization, which integrates the optimal evasion for linear cases as the baseline and trains to optimize the performance for nonlinear engagement with multiple pursuers. Therefore, offline training guarantees improvement of the constructed evasion guidance law over conventional ones. Ultimately, the guidance law for online implementation includes only analytical calculations. It maps from the confrontation state to the expected angle of attack and the unfolded angle while retaining high computational efficiency. Simulations show that the proposed evasion guidance law can utilize the change of unfolded angle to extend the maximum overload capability. And it surpasses conventional maneuver strategies by ensuring better evasion efficacy and higher energy efficiency.

Longitudinal Air-Breathing Hypersonic Vehicle Nonlinear Dynamic Simulation with Different Control Inputs

<div>The air-breathing hypersonic vehicle (AHV) holds the potential to revolutionize global travel, enabling rapid transportation to low-Earth orbit and even space within the next few decades. This study focuses on investigating the nonlinear dynamic simulation, trim, and stability analysis of a three-degrees-of-freedom (3DOF) longitudinal model of a generic AHV for variable control surface deflection, <i>δ<sub>e</sub></i> and <i>δ<sub>r</sub></i>. A simulation is developed to analyze the burstiness of the AHV’s nonlinear longitudinal behavior, considering the complete flight envelope across a wide range of Mach numbers, from <i>M</i> = 0 to 24, for selected stable <i>M</i>. The presented simulation assesses trim analysis and explores the dynamic stability of the AHV through its flight envelope and bifurcation method analysis is carried out to gain insight and validate the dynamic stability using eigen value approach.</div>

Numerical study for heat dissipation performance of phase transition cooling structure towards the wing antenna of hypersonic vehicle

Most airborne antennas of hypersonic vehicles rely on passive thermal protection, which is greatly dependent on material advancements. This paper introduces a wing antenna concept based on active cooling to enhance the high-temperature resilience of antenna materials for airborne applications. Developed an active cooling numerical model employing the volume of fluid (VOF) and coupling evaporative heat transfer, exploring the active cooling performance of the wing antenna unit construction using water as the coolant. Taking into account multiple structural parameters, establish a comprehensive performance factor as the optimization objective. The research analyzes the model's wall temperature distribution and the efficiency of active cooling. Under the same thermal conditions, the heat dissipation performance of the wing antenna unit varies when optimizing for heat transfer, pressure drop, and overall performance, respectively. The micro sized wing antenna unit optimized for heat transfer performance exhibited a wall temperature of just 325.8 °C under a heat flux of 1000 KW/m2, albeit with a notable pressure drop. Conversely, the unit optimized for comprehensive performance registered a wall temperature of 348.4 °C without a significant pressure drop. Established heat tests to validate the credibility of the numerical model. This paper provides several references for the design of phase-change heat dissipation structures and active cooling techniques for hypersonic vehicle wing antenna applications.

Supersonic cooling film flow evolution on a curved wall under hypersonic flow

Curvature plays a crucial role in evolving supersonic cooling film flow-field structures. Flow-field structural images were captured using nanotracer-based planar laser scattering,, and wall pressure values were obtained using experimentally validated numerical simulations. A supersonic cooling film is tangentially injected at the Mach number of Maj = 2.3 into a laminar boundary layer at a mainstream of Ma = 6. The supersonic cooling film inhibits mixing-layer instability on the convex curved wall (CV) and promotes it on the concave curved wall (CC). After increasing the total incoming pressure, the reduction ratio of static pressure (RSP) between the supersonic cooling film and the mainstream flow causes a delay in the position of the mixing-layer instability, smaller-scale vortex structures, and decreased flow velocity of the typical vortex structures on the CC and CV. The wall pressure increases for the CV and decreases for the CC, indicating that the supersonic cooling film suppresses the changes in wall pressure due to curvature. The supersonic cooling film suppresses the decrease in the impulses for bulk dilatation (Ip) due to convex curvature and the increase in Ip due to concave curvature. The growth rate of Ip on the CC increases from −15% to −8% and decreases on the CV from 31% to 12% in the bending impulse (IΦ) range of |IΦ| = 1.337–3.624 for a total inlet pressure of 0.5 MPa. Increasing the RSP could control the Ip values on curved surfaces more effectively. The results of this study can be applied to cooling the infrared optics window on hypersonic vehicles.

Upgraded VKI Plasmatron capabilities with supersonic nozzles for extended material characterization methods

In the domain of re-entry and hypersonic vehicles, advancing material development requires comprehensive material testing to thoroughly understand the aero-thermo-chemistry at the surface. This paper overviews hypersonic, high-temperature activities conducted at the von Karman Institute for Fluid Dynamics (VKI) Plasmatron facility, employing three recently developed supersonic nozzles. These activities aim to augment VKI’s experimental and numerical capabilities in high-temperature material research, contributing to a deeper understanding of the chemical surface interactions, at stagnation and off-stagnation locations, during re-entry. The newly developed nozzles were 3D printed using an additive manufacturing aluminum alloy. The three nozzles serve distinct purposes: two conical ones of varying lengths are designed for stagnation point testing, while the third nozzle, adopting a semi-elliptical shape, is tailored for off-stagnation point testing. The initial section of the paper focuses on designing, manufacturing, and simulating these nozzles, then we characterize the nozzle flow and compare with simulations, while the subsequent part will delve into showcases of recent activities conducted in the Plasmatron facility utilizing these distinct nozzle configurations.

Integrated entry guidance with no-fly zone constraint using reinforcement learning and predictor-corrector technique

This paper presents an integrated entry guidance law for hypersonic glide vehicles with no-fly zone constraint. Existing methods that employ predictor-corrector technique and lateral guidance logic for both guidance and avoidance, may have limitations in response time and maneuverability when facing sudden threats, because the guidance cycle is limited by computational efficiency and the bank angle magnitude cannot be adjusted according to the urgency of the avoidance. To overcome these challenges, the proposed method divides the entry process into safe flight stages and no-fly zone avoidance stages, and introduces reinforcement learning to develop an intelligent avoidance strategy for the latter. This division reduces the complexity of the learning problem by restricting the state space and increases the applicability in the presence of multiple no-fly zones. The trained avoidance strategy can directly output continuous bank angle command through a single forward calculation, considering both guidance and avoidance requirements. This enables the full utilization of the vehicle’s maneuverability and supports a high command update frequency to effectively handle threats. Additionally, a network trained via supervised learning is employed to generate reference commands, accelerating the training convergence of reinforcement learning. Simulation results demonstrate the effectiveness of the proposed guidance law, highlighting its high computational efficiency, command stability, and robustness. Importantly, the approach offers convenience in extending to multiple no-fly zones and accommodating vast initial state spaces.

Improvement and application of local-variable-based transition models to complex three-dimensional configurations

The hypersonic local-correlation-based transition modeling (LCTM) and amplification factor transport (AFT) transition models are extended to crossflow-induced transitions and applied to three-dimensional hypersonic vehicles. A new crossflow Reynolds number correlation is proposed and implanted to an existing hypersonic LCTM to extend it to take account of the crossflow effect. In the hypersonic AFT model, a new transport equation is added to predict the transported crossflow amplification factor. The two improved transition models are first validated using the HIFiRE-5 configuration. The results demonstrate that both transition models predict the transition on HIFiRE-5 with reasonable accuracy. Next, the two models are applied to the X-51A forebody and the X-33 vehicle. The results show that both models are capable of predicting transition behaviors and aeroheating environments on hypersonic complex geometries. The two models also provide reasonable reflections of the effects of various factors that influence transition, including the freestream Reynolds number, turbulence level, and angle of attack.

Quantification of plasma enabled surface cooling by electron emission from high temperature materials

In this work, the potential for hypersonic leading edge cooling by electron emission is demonstrated. To overcome space charge limitations the experiments are carried out in an argon discharge at 1 Torr. Cooling is observed with time-resolved measurements of the electron emission current and surface temperature, taking advantage of well controlled laser heating of the emitting surface and time accurate surface pyrometry. For the ignited mode of the plasma discharge, surface cooling by the electron emission is directly observed, leading to an estimated cooling rate of 1.6±0.2 MW m−2. Higher cooling rates with self sustained plasmas for space charge mitigation are expected using cesium transpiration. A two-dimensional model of heat transfer has been developed, which reproduces well the experimentally observed cooling dynamics. Parametric tests of emitter materials with various work functions show that for effective surface cooling by electron emission, the optimal work function must be less than 3.0 eV. This result indicates that electron cooling can be a promising thermal protection method for leading edges of hypersonic vehicles in flight.

Charging Process in Dusty Plasma of Large-Size Dust Particles

During reentry, the high temperatures experienced by near-space hypersonic vehicles result in surface ablation, generating ablative particles. These particles become part of a plasma, commonly referred to as a “dusty plasma sheath” in radar remote sensing. The dusty plasma model, integral in radar studies, involves extensive charge and dynamic interactions among dust particles. Previous derivations assumed that the dust particle radius significantly surpassed the Debye radius, leading to the neglect of dust radius effects. This study, however, explores scenarios where the dust particle radius is not markedly smaller than the Debye radius, thereby deducing the charging process of dusty plasma. The derived equations encompass the Debye radius, charging process, surface potential, and charging frequency, particularly considering larger dust particle radii. Comparative analysis of the dusty plasma model, both before and after modification, reveals improvements when dust particles approach or exceed the Debye length. In essence, our study provides essential equations for understanding dusty plasma under realistic conditions, offering potential advancements in predicting electromagnetic properties and behaviors, especially in scenarios where dust particles closely align with or surpass the Debye radius.

Influence of wall curvature on evolution of converged flow generated by crossing shock wave/boundary layer interaction: An experimental study

The curved surfaces widely exist in three-dimensional inward-turning inlets and variable sectional isolators, which determine the downstream evolution of the converged flow generated by the crossing shock wave/boundary layer interaction, thereby affecting the start performance and combustion efficiency of the hypersonic vehicles. In this work, the converged flow subjected to the curved surfaces is experimentally studied with the nanoparticle-based planar laser scattering method and particle image velocimetry. Results show that the velocity profiles of the converged flow on the transverse plane can be normalized at different angles. The upstream separation size is found to be decreased by the downstream convex surface due to the reduction of backpressure, especially in large-radius cases, whereas the converged flow gradually becomes smooth ought to the dependence of the acceleration effect on the velocity. The concave surface shows an opposite effect on the converged flow, which further indicates that a downstream convex surface is preferred for hypersonic inlets to weakens the flow non-uniformity to improve the inlet performance.

Performance comparison between closed-Brayton-cycle power generation systems using supercritical carbon dioxide and helium–xenon mixture at ultra-high turbine inlet temperatures on hypersonic vehicles

Closed-Brayton-cycle (CBC) has considerable potential for power generation on hypersonic vehicles, but its power level is limited by the cold source onboard, commonly the hydrocarbon fuel. To achieve power enhancement of CBC with finite cold source, this study assessed and compared the power output and thermal efficiency of supercritical carbon dioxide (S–CO2) and helium-xenon mixture (He–Xe) CBCs at ultra-high turbine inlet temperature (TIT), taking into account the thermal cracking of hydrocarbon fuel. Results indicate that raising TIT can significantly increase the electric power output for both CBC working fluids. The electric power under per unit mass flowrate of fuel is as high as 1600 kJ/kg at the TIT of 2000 K, making it easy to achieve megawatt-level power generation. The power enhancement of He–Xe CBC is more significant than S–CO2 CBC when the TIT exceeds 2000 K. Moreover, the compressor inlet temperatures and cycle pressure of S–CO2 CBCs are much higher than those of He–Xe CBC, causing additional challenges in the aerodynamic design and component manufacture. Therefore, He–Xe CBC operating at ultra-high TITs may be more suitable than S–CO2 CBCs for high-power electricity generation mission on hypersonic vehicles.

Intelligent maneuver strategy for hypersonic vehicles in three-player pursuit-evasion games via deep reinforcement learning.

Aiming at the rapid development of anti-hypersonic collaborative interception technology, this paper designs an intelligent maneuver strategy of hypersonic vehicles (HV) based on deep reinforcement learning (DRL) to evade the collaborative interception by two interceptors. Under the meticulously designed collaborative interception strategy, the uncertainty and difficulty of evasion are significantly increased and the opportunity for maneuvers is further compressed. This paper, accordingly, selects the twin delayed deep deterministic gradient (TD3) strategy acting on the continuous action space and makes targeted improvements combining deep neural networks to grasp the maneuver strategy and achieve successful evasion. Focusing on the time-coordinated interception strategy of two interceptors, the three-player pursuit and evasion (PE) problem is modeled as the Markov decision process, and the double training strategy is proposed to juggle both interceptors. In reward functions of the training process, the energy saving factor is set to achieve the trade-off between miss distance and energy consumption. In addition, the regression neural network is introduced into the deep neural network of TD3 to enhance intelligent maneuver strategies' generalization. Finally, numerical simulations are conducted to verify that the improved TD3 algorithm can effectively evade the collaborative interception of two interceptors under tough situations, and the improvements of the algorithm in terms of convergence speed, generalization, and energy-saving effect are verified.

Study of the thermal/thermodynamic performance of a bionic flat heat pipe designed to integrate thermal management and bearing capacity

When a hypersonic vehicle is flying at high speed, its rudder wing is subjected to severe aerodynamic thermal loads. In this paper, to meet thermal protection and load-bearing requirements during flight, a novel structure that integrates thermal management and carrying capacity is designed. A novel bionic flat heat pipe structure for integrating thermal management and bearing capacity requirements is designed using a beetle sheath wing as a biological template, and the thermal and mechanical performance of the structure is analysed. The interior consists of a 6061 aluminium bionic core covered with a nickel foam wick, and water is used as the working fluid. Ground tests are performed to verify the accuracy of the simulation results. The results show that the bionic flat heat pipe structure has a superior isotherm and bearing capacity, and the experimental data and finite analysis model results are in good agreement.

Aerodynamic Exergy-Based Analysis and Optimization of the Generic Hypersonic Vehicle Using FUN3D

The development of future novel aircraft concepts requires a holistic approach to vehicle design, analysis, and optimization. Aircraft designers can no longer consider components individually as systems become more interconnected and multidisciplinary. Hence, a new approach to aircraft design and a new way to measure the performance of aircraft are needed. A novel approach to aircraft design is the use of exergy methods that can evaluate typically disparate systems using a universal measure of performance mapped to global system performance. This paper introduces a new functional and its adjoint gradient in FUN3D to determine aerodynamic exergy destruction rates. The functional is verified using the Oswatitsch relationship by comparing it to native, surface-based drag coefficients. The adjoint gradient is verified using the FUN3D native complex step method, and discrete agreement is demonstrated for flowfields and geometric derivatives. The new functional is then used for aerodynamic exergy-based analysis and optimization of the generic hypersonic vehicle. An inverse design problem is conducted first to verify the design optimization framework. Finally, a planform and airfoil aerodynamic inviscid exergy optimization of the GHV is conducted, improving exergy destruction rates by 7.1% while making substantial improvements to the cruise trim characteristics of the vehicle.

Trajectory optimization of hypersonic vehicle considering the quasi-static assumption of pitch motion

The re-entry flight velocity of hypersonic vehicles is extremely high. If the trajectory precision can be improved in the optimization phase, the burden on the flight control system can be effectively reduced. Based on the phenomenon that the velocity of vehicle moving around the center of mass is small during the re-entry, a quasi-static assumption of the pitching motion for the re-entry vehicle is proposed. That is, the instantaneous balance of its pitching torque is assumed, and the torque balance is taken as the path constraint. Thus, a trajectory optimization method is established while considering the quasi-static assumption of pitching motion. When the minimum heating rate of the stagnation point or the smoothest trajectory is used as the optimization objective, respectively. Compared to traditional trajectory optimization methods that treat the vehicle as a mass point, the trajectory optimization method considers the quasi-static assumption of pitching motion, which can improve trajectory accuracy by an order of magnitude. In addition, solving the trajectory takes less than twice as long. This method can provide a reference for the offline/online trajectory design of re-entry vehicles.