Design Of Hypersonic Vehicles Research Articles

A New Very High-Order Finite-Difference Method for Linear Stability Analysis and Bi-orthogonal Decomposition of Hypersonic Boundary Layer Flow

Precisely predicting laminar-turbulence transition locations is essential for improvements in hypersonic vehicle design related to flow control and heat protection. Currently state-of-the-art eN prediction method requires the evaluation of discrete normal modes F and S for the growth rate of instability wave. Meanwhile, in receptivity studies, both the discrete and continuous modes, including acoustic, entropy, and vorticity modes, contribute to the generation of the initial disturbance. The purpose of this paper is to introduce a new very high-order numerical method to accurately compute these normal modes with finite-difference on a non-uniform grid. Currently, numerical methods to obtain these normal modes include two major approaches, the boundary value problem approach and the initial value problem approach. The boundary value approach used by Malik (1990) deploys fourth-order finite difference and spectral collocation methods to solve a boundary value problem for linear stability theory (LST). Nonetheless, Malik's presentation only demonstrated the computation of discrete modes, but not the continuous modes essential for conducting modal analysis on receptivity data. To obtain the continuous spectrum for his multimode decomposition framework, Tumin (2007) relies on an initial value approach based on the Runge Kutta scheme with the Gram-Schmidt orthonormalization. However, the initial value approach is a local method that does not give a global evaluation of the eigenvalue spectra of discrete modes. Furthermore, Gram-Schmidt orthonormalization, which can be error-prone in implementation, is required at every step of the integration to minimize the accumulation of numerical errors. To overcome the drawbacks of these two approaches, this paper improves the boundary value approach by introducing a new general very high-order finite difference method for both discrete and continuous modes eigenfunctions. This general high-order finite difference method is based on a non-uniform grid method proposed by Zhong and Tatineni (2003). Under the finite difference framework, discrete and continuous modes can be obtained by imposing proper freestream asymptotic boundary conditions based on the freestream fundamental solution behavior. This asymptotic boundary condition is used for obtaining both discrete and continuous modes that have both distinct (acoustic) and similar (vorticity and entropy) eigenvalues. Extensive verification of the new method has been carried out by comparing the computed discrete and continuous modes. Subsequently, the discrete and continuous modes obtained with this finite difference method are essential for the bi-orthogonal decomposition, which holds promising potential in obtaining an accurate evaluation of receptivity coefficients. The result of the bi-orthogonal decomposition for a hypersonic boundary layer flow over a flat plate is verified by comparing with existing results. Ultimately, the bi-orthogonal decomposition using the eigenfunctions has been applied to a case of freestream receptivity simulation for an axis-symmetric hypersonic flow over a blunt nose cone with modal contributions computed as coefficients for receptivity analysis.

Aerothermal and aerodynamic characteristics of reusable hypersonic vehicles with heat transfer minimised sweepback

Abstract Establishing aerothermal criteria for swept leading-edge hypersonic vehicle design is the predominant purpose of this work. This study is focused on two different configurations of vehicles based on the swept-back angle ( $\varLambda$ ) viz. minimum drag ( $\varLambda$ Drag-min ), and minimum heat transfer to vehicle ( $\varLambda$ HT-min ). Maximum wall temperatures obtained from the simulation performed in ANSYS 2020 with the k-epsilon turbulence model are 1,013 and 970K for $\varLambda$ Drag-min and $\varLambda$ HT-min , respectively. These temperatures are used to obtain the corresponding thicknesses of thermal protection systems to maintain inner wall temperature at 323K. Further study is divided into two cases depending on the direction of thickness of thermal protection system with respect to vehicle body. For constant payload capacity, the direction of thickness is outside; whereas for constant overall volume case, direction of thickness is inside. For constant payload volume case, the percentage weight reduction of thermal protection system is 4.8%. For constant overall volume case, the percentage payload capacity increases with design at $\varLambda$ HT-min by 4.04% in addition to thermal protection system weight reduction. The lift-induced drag on vehicles with design at $\varLambda$ HT-min is significantly reduced for both cases, by 47.68% (for constant payload volume) and 45.27% (for constant overall volume).

On the shock wave boundary layer interaction in slightly rarefied gas

The shock wave and boundary layer interaction (SWBLI) plays an important role in the design of hypersonic vehicles. However, discrepancies between the numerical results of high-temperature gas dynamics and experiment data have not been fully addressed. It is believed that the rarefaction effects are important in SWBLI, but the systematic analysis of the temperature-jump boundary conditions and the role of translational/rotational/vibrational heat conductivities are lacking. In this paper, we derive the three-temperature Navier–Stokes–Fourier (NSF) equations from the gas kinetic theory, with special attention paid to the components of heat conductivity. With proper temperature-jump boundary conditions, we simulate the SWBLI in the double cone experiment. Our numerical results show that, when the three heat conductivities are properly recovered, the NSF equations can capture the position and peak value of the surface heat flux, in both low- and high-enthalpy inflow conditions. Moreover, the separation bubble induced by the separated shock and the reattachment point induced by impact between transmitted shock and boundary layer are found to agree with the experimental measurement.

Experimental study of the effects of isolated roughness elements on the stability and transition of a hypersonic boundary layer on a flat plate

Surface roughness elements on hypersonic vehicles can cause early boundary layer transition, increasing wall skin friction and heat flux and affecting aircraft range and thermal protection systems. Accurate prediction of the transition caused by these roughness elements is crucial for the design of hypersonic vehicles. In this work, wind tunnel experiments on isolated roughness-induced boundary layer transition at Ma = 6 are conducted. Infrared thermography and high-frequency pressure sensors are utilized to investigate the effects of different roughness element configurations (cylindrical, diamond, ramp) on the hypersonic boundary layer instability and transition. The experimental results show that all three roughness elements can effectively enhance the generation of second mode waves and promote boundary layer transition. Compared to smooth surfaces, they exhibit similar frequency band range, faster growth, and earlier saturation. Among them, the ramp roughness element most effectively triggers the boundary layer transition, with a relatively small heat flux increase. Furthermore, bispectral analysis illustrates that all three roughness elements undergo self-interactions that lead to spectral broadening, ultimately resulting in boundary layer transitions.

Real Gas Effects on Receptivity to Roughness in Hypersonic Swept Blunt Flat-Plate Boundary Layers

Temperatures within the boundary layers of high-enthalpy hypersonic flows can soar to thousands or even tens of thousands of degrees, leading to significant real gas phenomena. Although there has been significant research on real gas effects on hypersonic boundary layer stability, their impact on the boundary layer’s receptive stage is still poorly understood. Most aerodynamic boundary layers in flight vehicles are three-dimensional. Because of complex geometry and significant crossflow effects, the crossflow mode in three-dimensional boundary layers is crucial in hypersonic vehicle design. In this study, a linear stability analysis (LST) accounting for chemical nonequilibrium effects (CNE) and its adjoint form (ALST) is developed to investigate the real gas effects on the stability and receptivity of stationary crossflow modes. The results indicate that real gas effects significantly influence the receptivity of stationary crossflow modes. Specifically, chemical nonequilibrium effects destabilize the crossflow modes but reduce the receptivity coefficients of the stationary crossflow modes. The Mach number effect was also investigated. It was found that increasing the Mach number stabilizes the stationary crossflow modes, but the receptivity coefficients increase. As the Mach number progressively rises, these effects alternately dominate, leading to a non-monotonic shift in the transition position.

Aerodynamic analysis of an airframe-inlet integrated full-waverider vehicle designed by osculating cone theory

Recently proposed, the airframe/inlet integrated full-waverider vehicle design methodology introduces an innovative approach to the design of air-breathing hypersonic vehicles. This methodology, compared to traditional integrated waverider design methods, leverages the features of the waverider to enhance the compression ability of the precompression surface and the lift-to-drag ratio characteristics of the vehicle. However, existing full-waverider design methods, based on cone-derived waverider theory, result in a gas cross-flow phenomenon at the inlet lip due to the conical shockwave. This paper suggests an airframe/inlet integrated full-waverider vehicle design method that utilizes the osculating cone waverider theory to improve gas intake quality. Initially, we present the method to obtain profile lines and basic flow field models at different osculating planes of the vehicle. We then validate the osculating cone integrated full-waverider vehicle design methodology through computational fluid dynamics. Although the viscous effect slightly reduces the lift-to-drag ratio of the vehicle and the total pressure recovery coefficient at the isolator exit, the shock wave remains attached to the leading-edge. Finally, a comparison with the cone-derived integrated full-waverider vehicle reveals that the osculating cone integrated full-waverider vehicle achieves similar aerodynamic performance, but with superior gas intake quality. The results indicate that the methodology proposed herein holds promise for guiding future engineering applications of integrated full-waverider vehicles.

Robust convex optimization for reentry glide trajectory using polynomial chaos

Aiming at the reentry glide trajectory design of hypersonic vehicles with uncertain parameters, a robust trajectory optimization method is proposed to enhance the anti-interference ability and process reliability. Firstly, a robust trajectory optimization model for re-entry gliding under uncertain conditions is constructed, which is transformed into a deterministic trajectory optimization problem with dimension expansion by using an uncertainty propagation algorithm based on the Gaussian quadrature and non-intrusive polynomial chaos. Then, the convex optimization technology is used to convexity and discretize the problem, and a deterministic trajectory optimization strategy based on the sequential convex optimization algorithm is designed to achieve a fast solution for the high-dimensional deterministic problem. Finally, the hypersonic glide reentry of the American X-33 is chosen as an example for the numerical simulation. The results show that comparing with the traditional deterministic trajectory optimization algorithm, the present method can effectively reduce the impact of random interference on the reentry glide trajectory and improve the reliability and robustness of the trajectory.

Cooperative game penetration guidance for multiple hypersonic vehicles under safety critical framework

The rapid development of the anti-missile weapon technology brings new challenges to the cooperative penetration strategy solution and the guidance law design for Hypersonic Vehicles (HVs). This paper studies the coordinated game penetration guidance problem for multiple hypersonic vehicles faced with space threat areas. A scheme for seeking cooperative game penetration guidance strategy under safety critical control framework is presented. In this scheme, a multi-HV cooperative game model is proposed in a minimum optimization form which can simplify the solving process and accelerate the computing speed. Then, a second-order control barrier function is developed to transform the implicit nonlinear constraints of the proposed model into linear ones. In order to obtain better performance of guidance strategy, a composite guidance law under the safety critical control framework is presented to allocate guidance strategies appropriately in the whole process. It is shown that the proposed scheme can guarantee successful penetration while avoiding threat areas. Finally, a comparative simulation with a two-on-three game is conducted to verify the effectiveness of the proposed method.

Snap-through response of three-dimensional braided composite panels in thermal-acoustic-aero coupled field

This paper presents a novel investigation into the snap-through response of three-dimensional (3D) four-directional braided composite panels subjected to aerodynamic loads generated by supersonic flow. The study introduces a unique finite element method to establish the dynamic equations of the panel. Utilizing a multi-scale approach, the thermo-acoustic and aerodynamic coupling response of simply supported 3D four-directional braided composite panels is analysed in both the time and frequency domains. The paper explains the underlying mechanism of snap-through motion using the potential well theory and introduces the concept of zero-crossing frequency for quantitative classification of snap-through motion. Notably, variations in the critical sound pressure level (SPL) due to changes in aerodynamic and acoustic loads are identified. The research reveals that when the dynamic pressure is below the critical flutter dynamic pressure, the potential well deepens, making it more difficult for panels to experience snap-through motion. Conversely, once the dynamic pressure surpasses the critical flutter dynamic pressure, persistent snap-through motion is triggered. Moreover, the aerodynamic load induces changes in the potential well, leading to a decrease in the snap-through critical SPL. These findings not only contribute to the prediction of fatigue life in braided composite panels but also have practical implications for the design of hypersonic vehicles.

Difference analysis in terahertz wave propagation in thermochemical nonequilibrium plasma sheath under different hypersonic vehicle shapes

Research on terahertz communication under different vehicles has important guiding significance for the design of future hypersonic vehicle and Radio Frequency (RF) blackout. In this paper, a joint simulation model of plasma flow under thermochemical nonequilibrium state and terahertz transmission is developed to investigate the differences in terahertz wave transmission characteristics under different vehicle shapes and the related mechanisms. By comparing the plasma sheath characteristics and terahertz transmission among HIFIRE-5b (Hypersonic International Flight Research Experimentation-5b), RAM C(Radio Attenuation Measurement C) and ARD (Atmospheric Reentry Demonstrator) vehicles, it is found that the sheath thickness and electron density of ARD vehicles is significantly larger than that of HIFIRE-5b and RAM C vehicles, resulting in greater terahertz wave attenuation. Collision absorption plays a major role in the terahertz attenuation of vehicles, and the contribution of reflection effects is only observed in the ARD vehicle due to its larger plasma sheath thickness and spatial structure variation. Based on the above comparison results, a shape design scheme of reducing the vehicle head and tail for mitigating RF blackout is proposed, and the scheme is proved by further analyzing the effects of different vehicle heads and tails on the terahertz communication. With the decrease of the head radius and tail width of hypersonic vehicle, the wave attenuation at the same terahertz frequency decreases, and the contribution of reflection effect to wave attenuation gradually disappears. Therefore, the shape design scheme of reducing the vehicle head and tail can effectively alleviate the RF blackout problem, which provides an important reference value for future hypersonic vehicle design.

Characterization of nonequilibrium shock interaction in CO2-N2 flows over double-wedges with respect to Mach number and geometry

The characterization of the shock interaction mechanism originating from the high-Mach nonequilibrium flow over double-wedge geometries is key to the design of hypersonic vehicles. The impact of changes in the freestream Mach number and double-wedge geometry on the patterns of shock interaction is investigated by means of numerical simulation in the case of CO2-N2 flows. The extended laminar Navier–Stokes equations with a two-temperature model to account for translational-to-vibrational internal energy transfer are considered the physical model of this type of flow. Simulations show that reducing the freestream Mach number leads to an increase in the separation region, both in the compression corner and in the locations of shock impingement. The impact of the size of the separation region on the patterns of interaction is such that it causes variations in the type of shock interaction. From the point of view of the flow physics near the wedges, decreasing the freestream Mach number has an equivalent effect to increasing the angle of the second wedge and an opposite effect to increasing the freestream temperature on the pattern of interaction. Results show that decreasing the freestream Mach number leads to an overall reduction in pressure and heating loads along the surface of the wedges and smaller regions of thermal equilibrium behind the bow shock.

Effect of Rarefaction on Thermal and Chemical Non-Equilibrium for Hypersonic Flow With Different Enthalpy and Catalytic Wall Conditions

AbstractCompressibility and rarefaction effect plays an essential role in the design and study of objects experiencing hypersonic flows. The presence of chemical and thermal non-equilibrium in hypersonic flows increases the complexity of estimating aerothermodynamic properties, which are essential for developing thermal protection systems and the aerothermodynamic design of hypersonic vehicles. In this study, the hy2Foam solver, developed in an OpenFOAM framework by hyStrath group, is used to understand the effect of Knudsen number (which in turn depends on the altitude) and freestream enthalpy variation on the surface aerothermodynamic properties such as pressure, heat flux, velocity slip, temperature jump, and flow field variables such as species concentration and temperature, in five-species air flow over a cylinder, for both noncatalytic and fully catalytic wall conditions. The novelty of the work lies in reporting the effect of rarefaction on thermal and chemical non-equilibrium (associated with hypersonic flows), and thus on the surface properties under different enthalpy and wall catalytic conditions. It has been shown that the rarefaction effect is more pronounced on the vibrational temperature component and for high enthalpy gas. The surface wall heat flux and the chemical reaction rate among the species decrease with rarefaction. The skin friction coefficient is one of the most sensitive properties, while the pressure coefficient has been the least susceptible to non-equilibrium effects. The stagnation points heat flux at different Knudsen numbers shows good agreement with the existing correlation in literature for both low and high enthalpy flows, which further establishes the validity of the study done in this work.

Dynamic Transition of Unsteady Supersonic Flow From Mach to Regular Reflection Over a Moving Wedge

Abstract The design of supersonic and hypersonic air-breathing vehicles is influenced by the transition between the Mach reflection (MR) and regular reflection (RR) phenomena. The purpose of this study is to investigate the dynamic transition of unsteady supersonic flow from MR to RR over a two-dimensional wedge numerically. The trailing edge of the wedge moves downstream along the x-direction with a velocity, V(t) at a freestream Mach number of 3. An unsteady compressible inviscid flow solver is used to simulate the phenomenon. The Arbitrary Lagrangian–Eulerian (ALE) technique is applied to deform the mesh during the wedge motion. The dynamic transition from MR to RR is defined by two criteria, the sonic and the von Neumann. Moreover, the lag in the dynamic transition from the steady-state condition is studied using various nondimensional angular velocities, κ, in the range of [0.1-2]. The lag effect in the shock system is remarkable at the high values of κ greater than 1.5. Furthermore, the dynamic transition from MR to RR happens below the dual solution domain (DSD) because the shock is upstream curved during the wedge motion.

An experimental study on the stability of hypersonic plate boundary layer regulated by a plasma actuation array

Laminar–turbulent transition control is one of the key problems in the design of hypersonic vehicles. In particular, for air-breathing hypersonic vehicles, early onset turbulence in the boundary layer of a scramjet inlet precursor is desirable. Plasma actuation has become a popular flow control method in recent years and has made progress in regulating the stability of supersonic boundary layers. However, there have been few studies on the regulation of the stability of hypersonic boundary layers. In this paper, an experimental study on the stability of the hypersonic plate boundary layer is carried out using a spanwise plasma actuation array. The characteristics and evolution of different kinds of unstable waves in the hypersonic plate boundary layer (especially the dominant first mode wave) are analyzed based on the results of linear stability theory and high-frequency pulsation sensors. The typical morphological characteristics of the boundary layer and the macro-control effect of the plasma actuation array are explored through a high-speed schlieren method. Finally, based on grayscale mode extraction and proper orthogonal decomposition, the influences of three different actuation frequencies on the instability waves and characteristic structure of the boundary layer are studied, including the dominant frequency of the first mode wave, the half-frequency of the dominant frequency of the first mode wave, and the dominant frequency of the second mode wave, the change of characteristic structures under the regulation of plasma actuation is further discussed. The corresponding regulation rules and mechanisms are summarized. The results show that the plasma actuation array can advance the starting position of laminar discontinuities and that the induced coherent structure can excite instabilities at an earlier flow-direction position. The actuation can be used to control the stability of the boundary layer by acting on the first mode wave to break the original unstable wave spectrum characteristics. This verifies the ability of extensional array plasma actuation to regulate the stability of the hypersonic plate boundary layer and suggests it has great potential in the promotion of hypersonic boundary layer transition.

Appointed-Time Control for Flexible Hypersonic Vehicles with Conditional Disturbance Negation

This paper investigates the anti-disturbance control design for flexible air-breathing hypersonic vehicles (FAHVs) with appointed-time prescribed tracking performances. The challenging issues include multisource disturbances, cross-coupling effects of the vehicle dynamics, and asymmetric amplitude and rate saturations. To address these issues, we propose an appointed-time prescribed performance control scheme for FAHVs via conditional disturbance negation technique. In contrast to existing disturbance rejection approaches, the proposed control scheme not only estimates the disturbances first in a fixed time, but also evaluates the estimated disturbances and selectively conducts compensation actions according to the insight of the FAHV dynamic characteristics. To further enhance convergence rate, the composite appointed-time prescribed performance controllers are designed for FAHVs via time-varying barrier Lyapunov function and nonsmooth backstepping technique, which ensure satisfactory transient response and steady-state performances. In addition, the asymmetric amplitude and rate saturation problem of actuators are dealt with by introducing unified approximation dynamics. It is rigorously proved that the practical appointed-time convergence of the tracking errors and the fixed-time convergence of all signals in the resultant FAHV closed-loop system can be achieved under the proposed control scheme. Finally, extensive comparative simulations are provided to demonstrate the feasibility and superiority of the proposed approach.

Effects of cavity shapes and sizes on rarefied hypersonic flows

Cavity structures are frequently encountered on hypersonic vehicle surfaces, which change the aerodynamic heat flux and force (pressure) compared with flat vehicle surfaces. The effects of the cavity length-to-depth ratio, cavity rounded corners, Maxwellian accommodation coefficient, and free-stream Knudsen number on the heat flux and pressure loaded on the vehicle surface with rectangular cavities have been studied extensively. Still, the influence of the cavity shape and size remains. In this study, the effects of cavity shape and size on the flow field characteristics and aerodynamic surface quantities on the surface are analyzed using the direct simulation Monte Carlo method. The rectangular cavities with sidewall angles (α) of 0°, 15°, 30°, and 45° are selected to investigate the shape influence, and the semicircular cavities with radii (R) of 3, 6, 9, and 12 mm are chosen for size effect. The results indicate that the gas flows over the cavity configurations specified in our study hold the “open-cavity” characteristics. The vortexes formed inside the cavity are shrunk as α increases from 0° and 45°, but they are similar under the R-range explored. The pressure, skin friction, and heat transfer coefficients over the surface display complex trends as varying α and R. In addition, the effect of cavities on the overall aerodynamic heat flux and force, which are vital for practical applications, are explored. The results reveal that the aerodynamic heat flux is dramatically influenced by α and R, while the aerodynamic force is mildly affected. Overall, the larger the α and R, the greater the effect of the cavity on the rarefied hypersonic flows. Our research is beneficial to the protection and optimal design of hypersonic vehicles.

Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Hypersonic vehicle design requires mitigation of the high heat fluxes present in regions of shock-wave/boundary-layer interactions. A candidate technology that may be applied locally to these regions is transpiration cooling. In this work, experiments were conducted in the University of Oxford’s high-density tunnel at Mach 6.1 in both laminar and turbulent undisturbed boundary-layer regimes where a 10 deg shock generator impinged a strong oblique shock wave onto a transpiration-cooled microporous injector. For the laminar boundary layer, due to the strength of the incident shock, a transitional shock-wave/boundary-layer interaction region was formed with peak heating over 50 times greater than the nominal laminar level. Both nitrogen and helium were used as coolants. Relatively low levels of helium injection of for the transitional and for the turbulent scenarios were sufficient to reduce the heat transfer downstream of shock interaction to approximately 50% of the value without cooling. In fact, helium is highly effective with a similar cooling performance achieved as eight times the equivalent mass flux of nitrogen. The experimental data are correlated, and both the turbulent and transitional shock-impingement scenarios display a similar trend of reduced surface heat transfer with higher blowing parameters. Empirical fits are proposed that may be used for initial systems design.

Heat Flux Identification of Aircraft Structure with Artificial Neural Network Compensation

Aerodynamic heating measurement of reusable hypersonic vehicles has always been an important aspect of hypersonic vehicle design. In this paper, a mechanistic-model-based heat flux identification method with artificial neural network (ANN) compensation is established to determine the spatially distributed heat flux of the aircraft structure. A one-dimensional heat conduction model is used to estimate heat flux by a robust and efficient algorithm integrating Tikhonov regularization with Levenberg–Marquardt method. The one-dimensional estimated heat flux has large errors for not considering multidimensional heat conduction effects. The proposed mechanistic-model-based method is then utilized to compensate the multidimensional heat conduction by ANN. The performance of the proposed method will be assessed by the determination of the heat flux of a two-dimensional plate and aircraft structure. Results show that compared with the one-dimensional inversion results, ANN compensation method can significantly improve the accuracy of estimated heat flux and is also applicable for larger levels of heat flux. The proposed compensation method is an effective technique to identify the nonuniform surface heat flux of multidimensional structures.

Constrained Integrated Guidance and Control Scheme for Strap-Down Hypersonic Flight Vehicles with Partial Measurement and Unmatched Uncertainties

This paper investigates the issue of integrated guidance and control (IGC) design for strap-down hypersonic flight vehicles with partial measurement information and unmatched uncertainties. A constrained IGC scheme is proposed by combining the barrier Lyapunov function-based backstepping methodology and the specific output-based finite-time disturbance observer. Different from the existing methods, which require the state information and matched disturbances, the main features of the presented approach is capable of addressing the partial measurement knowledge and unmatched uncertainties simultaneously. The IGC model of hypersonic flight vehicles is first formulated, and based on that, the specific output-based finite-time disturbance observer (OFTDO) is proposed to achieve the finite-time estimation of the unmatched uncertainties through the output. Then, the constrained IGC strategy is constructed via the backstepping technique. The stability of the closed-loop system including the estimation and tracking errors dynamics is analyzed in detail. The effectiveness of the proposed method is verified by numerical simulations and Monte-Carlo tests.

Aerodisk Effect on Hypersonic Boundary Layer Transition and Heat Transfer of HIFiRE-5 Vehicle

The substantial aerodynamic drag and severe aerothermal loads, which are closely related to boundary layer transition, challenge the design of hypersonic vehicles and could be relieved by active methods aimed at drag and heat flux reduction, such as aerodisk. However, the research of aerodisk effects on transitional flows is still not abundant. Based on the improved k-ω-γ transition model, this study investigates the influence of the aerodisk with various lengths on hypersonic boundary layer transition and surface heat flux distribution over HIFiRE-5 configuration under various angles of attack. Certain meaningful analysis and results are obtained: (i) The existence of aerodisk is found to directly trigger separation-induced transition, moving the transition onset near the centerline upstream and widening the transition region; (ii) The maximum wall heat flux could be effectively reduced by aerodisk up to 52.1% and the maximum surface pressure can even be reduced up to 80.4%. The transition shapes are identical, while the variety of growth rates of intermittency are non-monotonous with the increase in aerodisk length. The dilation of region with high heat flux boundary layer is regarded as an inevitable compromise to reducing maximum heat flux and maximum surface pressure. (iii) With the angle of attack rising, first, the transition is postponed and subsequently advanced on the windward surface, which is in contrast to the continuously extending transition region on the leeward surface. This numerical study aims to explore the effects of aerodisk on hypersonic boundary layer transition, enrich the study of hypersonic flow field characteristics and active thermal protection system considering realistic boundary layer transition, and provide references for the excogitation and utilization of hypersonic vehicle aerodisk.