The High-pressure Capturing Wing (abbreviated as HCW) aerodynamic configuration significantly enhances the vehicle's lift and lift-to-drag ratio by effectively utilizing the strong shock-dominated flow characteristics under hypersonic conditions. According to its design theory, the optimal position of the capturing wing (abbreviated as CW) is directly correlated with the freestream Mach number. Addressing the wide-Mach-number flight characteristics of hypersonic glide vehicle (abbreviated as HGV), this study proposes a novel concept of morphing HCW configuration, composed of an axisymmetric fuselage with the CW. The design principle for the morphing configuration was investigated under a typical glide trajectory with considerations of the effects of both Mach number and viscosity. Based on the design principle, a morphing configuration was developed, and aerodynamic characteristic analysis was subsequently conducted. The results indicate that as the freestream Mach number increases, the shock layer gradually becomes thinner. Meanwhile, the wall boundary layer thickness gradually grows due to the enhanced viscosity effect. The coupled influence of Mach number and viscosity causes the fuselage shock angle to first decrease and then increase. These flow characteristics determine that the horizontal position of the CW gradually shifts rearward, while the vertical position first descends and then ascends as the Mach number increases in the morphing configuration design. The morphing configuration yields relatively small displacement of the CW’s position and significant aerodynamic improvements under off-design conditions, with the CW’s lift augmentation reaching 21%-33% and the vehicle’s lift-to-drag ratio enhancement achieving 16%-27%. Additionally, the aircraft's center of pressure exhibits a reduction in its variation with increasing Mach number.

The aerodynamic performance of the high-pressure capturing wing(HCW),a novel aerodynamic configuration with significant promise for high-speed aircraft,primarily depends on its geometry and positioning.This configuration differs significantly from conventional ones,as the severe shock wave generated by the airframe strongly interferes with the HCW at high speeds.To investigate the influence of airfoil geometry on the aerodynamic characteristics of this configuration,this study adopts a single-wing HCW concept and selects four typical high-speed airfoils(biconvex,quadrilateral,hexagonal,and triangular)to analyze the effects of airfoil shape on lift-to-drag characteristics and stability under design(freestream Mach number 6)and off-design(freestream Mach number 3)conditions via numerical simulations.The results show that under the design condition,the variation in the overall lift-to-drag ratio caused by different airfoils is relatively small,not exceeding 2.2%,while under the off-design condition it increases to 4.9%.This difference is attributed to the interference of the leading-edge shock wave and reflected shock wave of the capturing wing with the airframe under the off-design condition.Among the four airfoils,the quadrilateral airfoil exhibits the highest overall lift-to-drag ratio under both operating conditions.Stability analysis indicates that,except for the triangular airfoil,for which the center of pressure and the aerodynamic center shift rearward by 3%-4%at most angles of attack,the differences among the other airfoils are small.Therefore,under the design condition and with the same maximum relative thickness,the aerodynamic characteristics of the HCW configuration are not particularly sensitive to the choice of airfoil(except for the stability characteristics of the triangular airfoil),which provides broader optimization space for wide-speed-range airfoil design.

In recent decades, wide-speed-range aircraft with hypersonic cruise capability have garnered significant attention and emerged as a prominent research focus worldwide. However, a wider flight velocity range poses significant challenges to aerodynamic configuration design. To effectively overcome the trade-off between lift-to-drag ratio and volumetric efficiency, we propose a novel aerodynamic configuration with high volumetric capacity based on the high-pressure capturing wing concept. Numerical simulations demonstrate that this configuration achieves a maximum lift-to-drag ratio of 5.89 at Mach number 6.0, representing an improvement of over 18% compared to a reference configuration without the high-pressure capturing wing. Furthermore, numerical simulations and analyses were conducted focusing on two typical transonic conditions at Mach number 0.8 and 1.2. The results demonstrate that compared to the reference configuration, the new high-pressure capturing wing configuration exhibits increased lift and drag coefficients to varying degrees under both transonic conditions. Although this comes with some degradation in lift-to-drag ratio, the shift of the aerodynamic center is significantly reduced across the wide speed range. Specifically, the relative shift of the aerodynamic center is reduced by 11.1% in the transonic regime and by 49.9% across the transonic-to-hypersonic regime. Further analysis reveals that this reduction in aerodynamic center shift is primarily attributed to the coupling effects between the additional wing surface and the fuselage.

Wide-speed-range hypersonic vehicles represent a major research focus in aerospace. However, their aerodynamic configuration design faces inherent challenges such as the contradiction between volumetric efficiency and hypersonic lift-to-drag ratio, as well as the conflict between high-speed and low-speed aerodynamic configurations. Facing these challenges, the I-plane configuration based on the high-pressure capturing wing (HCW) principle demonstrates considerable potential. It provides simultaneously high volumetric efficiency, high lift, and high lift-to-drag ratio under hypersonic conditions, while its dual lifting surfaces enhance aerodynamic performance in the subsonic regime. In particular, the incidence angle of the independent HCW is a key design parameter influencing the vehicle's wide-speed-range aerodynamics. Therefore, this paper systematically examines the effects of varying the HCW incidence angle on the aerodynamic characteristics of a novel I-plane configuration under subsonic and hypersonic conditions, using Computational Fluid Dynamics (CFD) numerical simulations and Radial Basis Function (RBF) surrogate models. The results show that under subsonic conditions, an increase in the incidence angle produces a marked rise in the overall lift. With the HCW set at a larger incidence angle, overall lift increases by >40% relative to the reference configuration without HCW. Under hypersonic conditions, an increase in the incidence angle causes the lift to drag ratio to first rise and then decline. At a smaller incidence angle, the maximum lift-to-drag ratio reaches 6.20, corresponding to a 22.5% improvement over the reference configuration. The optimal HCW incidence angle varies for different speed ranges. Leveraging the HCW's structural separation from the fuselage, a dedicated variable configuration design can be implemented for the HCW itself, thereby enhancing wide-speed-range aerodynamic performance.

The High-pressure Capturing Wing aerodynamic configuration (abbreviated as HCW) can enhance the aircraft's lift and lift-to-drag ratio through the effective utilization of favourable aerodynamic interference between the fuselage and the capturing wing (abbreviated as CW). Consequently, the fuselage shape significantly influences not only the aerodynamic characteristics of the fuselage but also those of the CW. In this study, based on the concept of a fuselage of revolution combined with CW, the effects of different fuselage profile shapes-specifically, conical profile, 3/4-power profile, and Karman profile-on the aerodynamic characteristics of the HCW configuration are systematically investigated using a numerical simulation technique. The results indicate that while the 3/4-power and Karman configurations differ from the conical configuration by employing non-uniform compression, which increases the fuselage volumes by 18.5 % and 48.1 %, respectively, they achieve drag reductions of 6.7 % and 1.1 % under the design condition at Ma = 6. Moreover, due to the influence of different fuselage profiles on the flow characteristics, the peak pressure on the lower surface of the CW is reduced for both the 3/4-power and Karman configurations, while the area of the high-pressure zones is increased. Overall, compared to the conical configuration, the 3/4-power configuration exhibits a slight decrease in lift, whereas the Karman configuration demonstrates a significant increase in lift. Both configurations achieve an increase in the lift-to-drag ratio. Additionally, an analysis of the effects of different Mach numbers (Ma = 7, Ma = 8) that as the Mach number increases, the lift becomes more sensitive for both the 3/4-power and Karman configurations. However, compared to the conical configuration, both configurations keep exhibiting a decrease in overall drag and an increase in the lift-to-drag ratio. Furthermore, under different Mach number conditions, the centers of pressure for both configurations shift rearward.

The influence of sideslip angle on the 3-D asymmetry swept-forward fin shock interactions are analyzed by numerical simulation under Mach number 6.36, Reynolds number 3.27 × 10 7 /m and wedge compression angle 12°. Firstly, the effect of different swept-forward angles on the flow field structure and aerodynamic thermal characteristics under 0° sideslip is analyzed. With the increase of the swept-forward angle, there are three kinds of shock wave interaction in turn on the symmetry plane: type IV interaction, the combination of type II and type III interaction, and type I interaction. Then, a typical swept-forward angle is selected for each interaction type, and the flow characteristics at 10° and 20° sideslip angles are simulated. The results indicate that: Under the influence of the sideslip, the heat distribution on the fin becomes asymmetric, and the maximum heat flux position at the fin’s leading edge shifts from the symmetry plane to the windward side. The sideslip angle affects the maximum heat flux with changes within 10%. At all swept-forward angles, the windward heat flux increases significantly on the sidewall near the fin’s leading edge as the sideslip angle increases. At a 20° sideslip angle, for swept-forward angles of 20° and 30°, the maximum heat flux on the sidewall near the fin’s leading edge exceeds the heat flux at the stagnation point of the two-dimensional cylinder with the same radius under free flow. For swept-forward angles of 45°, the maximum heat flux on the sidewall near the leading edge is 0.74 times that of the two-dimensional cylindrical stagnation heat flux of the same radius under free flow.

Under beneficial aerodynamic interference, the innovative high-pressure capturing wing (HCW) configuration exhibits remarkable aerodynamic performance at hypersonic speeds. At the same time, the additional lifting surface of the HCW may significantly improve the lift at subsonic speeds, which makes the configuration a promising concept for air vehicle design covering a wide speed range. Recent research has shown that the HCW can generate flow separations on the upper surface of the fuselage in certain subsonic flow conditions, which considerably deteriorates aerodynamic performance of the air vehicle. In this paper, a hybrid configuration combining HCW with a cone-truncated cone fuselage is studied. Numerical simulations are conducted to examine the impact of fuselage corner bluntness on flow characteristics and aerodynamic performances in typical subsonic (Mach 0.7) and hypersonic (Mach 7) flow conditions. The results show that at Mach 0.7, with the increase of blunt radius at the fuselage corner, the flow separations on the upper surface of the fuselage can be effectively restrained. The lift coefficient of the vehicle remains virtually unchanged, while the drag coefficient decreases significantly by 69% when the blunt radius is 600 mm. At Mach 7, the fuselage corner bluntness leads to a slight drop of 9% and 8% in the lift and drag coefficient of the vehicle, respectively, while the lift-to-drag ratio remains virtually unchanged.

The transition from type I to II in the swept-forward fin–shock interaction is studied both numerically and analytically. The transition conditions are numerically obtained using an inviscid computational fluid dynamics method. The results are compared with the classical three-shock theory, in which the detachment criteria are computed based on the parameters immediately after the shock. A discrepancy is observed, where the critical swept-forward angles from the numerical results are always larger, with a maximum deviation of 5°. Detailed flow patterns indicate this deviation is caused by the downstream of the bow shock wave. By analyzing the shock patterns in a wide range of both the inclined shock angles and the swept-forward angles, three mechanisms are identified. The first mechanism is related to the Mach reflection formed by a transmitted shock reflecting on the wall. When the top of the Mach stem reaches the fin–shock interacting point, the interaction transits from type I to II. The second is associated with a secondary transmitted shock wave achieving its maximum deflection angle. The last is due to the subsonic channel linking the symmetric and lateral planes. The quantities of the advanced transition are derived theoretically and in good agreement with the numerical results. The results have potential to benefit the aerodynamics and aerothermodynamics design of hypersonic vehicles.

With an additional wing upon the fuselage, the novel high-pressure capturing wing (HCW) configuration exhibits remarkable aerodynamic characteristics at hypersonic speeds under beneficial aerodynamic interference. This bi-wing structure can also enhance the lift at subsonic speeds, positioning HCW configuration as an excellent aerodynamic layout under wide-speed range conditions. In this paper, a single-wing principle HCW configuration is carried out to analyze the influence of the variations in the geometric parameters of the HCW on the aerodynamic performance. Drawing on extant research, four key geometric parameters of the HCW are chosen as design variables, and the hypersonic as well as supersonic conditions are selected for surrogate-based shape optimization. Utilizing polynomial response surface method and method of moving asymptotes, the single-objective optimization studies are carried out with the objectives of maximum lift-to-drag ratio at Ma = 6 and minimum drag coefficient at Ma = 3. Subsequently, the sensitivity analysis is performed for each design parameter. The above methods exhibit notable precision and favorable results. The results show that except for the leading-edge sweep angle, the other three optimization results exhibit divergent trends in their variations. The setting angle has the most significant influence on the aerodynamic forces, owing to the influences of its variation on the shock wave intensity and reflection angle. Based on the stronger intensity of shock wave, this sensitivity indices are higher at Ma = 6. The other three parameters (half-span, leading-edge sweep angle, and trailing-edge sweep angle) modify the aerodynamic forces by adjusting the area of high-pressure region on HCW. Due to the different flow field structures, optimum parameters exhibit diverse effects on lift coefficient and lift-to-drag ratio.

The geometrical properties of streamlines, such as the curvatures, directions and positions, are studied in steady inviscid compressible flow fields via differential geometry theories and conservation laws. The influences of the streamline geometries on the flow speeds and pressures are also identified and discussed. By transforming the streamlines to fill the domain and satisfy the boundary conditions, a unified geometry-based solver, the streamline transformation method, is proposed for both subsonic and supersonic regions. The governing equations and boundary conditions along streamlines and shock waves are also derived. This method is verified by numerical results of three typical flow fields, including the subsonic channel flow, the supersonic downstream of attached shock waves and especially the subsonic/supersonic downstream of detached bow shock waves. Both two-dimensional planar and axisymmetric flow fields are considered. Compared with the results from computational fluid dynamics, good agreements are achieved by this method, while fewer computational resources, by an order of magnitude, are consumed. Features of these flow fields are also analysed from a geometrical perspective, such as flow speeds and pressures deviated by the wall curvatures, and three-dimensional effects in the after-shock flow fields. For a hyperbolic-shaped bow shock wave, the stand-off distances and the transitions from subsonic to supersonic regions are also discussed. As indicated by the accuracy, efficiency and applicability in a wide range of flow speeds, the streamline transformation method would be a potential candidate for the theoretical analysis and inverse design of high-speed flow fields, especially where the subsonic regions exist downstream of strong shock waves.

A hybrid approach based on the immersed boundary method (IBM) is developed for computation of flow-induced sound around moving bodies. In this method, a high-fidelity direct numerical simulation (DNS) solver is used to simulate the incompressible flow field. The sound field is predicted by discretizing acoustic perturbation equations (APEs) with dispersion-relation-preserving space scheme and low-dispersion and low-dissipation Runge-Kutta time integration. A sharp-interface IBM based on ghost-cell is implemented for present two-step DNS-APE approach to deal with complex moving bodies with Cartesian grids. The present method is validated through simulations of sound generation caused by flow past a rotating cylinder, an oscillating cylinder, and tandem oscillating and stationary cylinders. The sound generated by typical kinds of complicated bio-inspired locomotions, i.e., flapping flight by wings of varied shapes and collective undulatory swimming in tandem, are investigated using present method. The results demonstrate potential of the hybrid approach in addressing flow-induced sound generation and propagation with complex moving boundaries in a fluid medium, especially for the sound characteristics of bio-mimetic flows, which might shed lights on investigations on bio-acoustics, ethology of complex animal system, and related bio-mimetic design for quietness.

纵向稳定性是关乎临近空间高超声速飞行器操稳特性、飞行安全和飞行品质的关键因素之一.目前纵向稳定性的研究主要围绕常规单升力面气动布局,而对具备双升力面结构的高压捕获翼新型气动布局,相关研究还有待加强.本文面向一种高压捕获翼飞行器,结合小振幅强迫振荡法、非定常数值模拟方法和最小二乘方法辨识纵向气动导数,详细研究了飞行姿态、振荡参数、来流条件和质心位置等因素的不确定性对飞行器纵向静、动稳定特性的影响规律.研究发现,当平衡攻角从-10°逐渐增大到14°时,纵向静稳定性整体呈现出波动减弱的趋势,而动稳定性却先小幅波动后迅速增强.振荡频率越大时,纵向稳定性越强,但总体影响不太明显;当振荡幅值低于2°时,气动导数的辨识结果相近,而在较大的振荡幅值下,最大攻角附近非定常气动力的迟滞效应显著增强.飞行马赫数增大时,静、动稳定性均显著减弱,且对较小的飞行马赫数比较敏感.飞行高度增大时,静稳定性逐渐减弱但动稳定性却逐渐增强,且均对较大的飞行高度更加敏感.飞行器纵向质心位置的改变主要对静稳定性影响较大,而对动稳定性的影响却相当有限.

Compressible flows typically exhibit multiple shock waves which interact with each other, making the detection of these shock waves crucial for various aspects of flow studies including construction of high-order numerical schemes (e.g., shock-fitting), adaptive grid refinement, and flow visualization. This study aims to effectively identify and localize multiple shock waves and their interaction points in two-dimensional inviscid steady and unsteady flows. A novel shock wave pattern recognition method based on cluster analysis is proposed, including three processes. First, a series of grid-cells located at the transition zones of captured shock waves are extracted using a shock wave detection approach based on local flow variation. Subsequently, these grid-cells are grouped into numerous clusters using the classical K-means clustering algorithm, with categorization based on nearest neighbor features. Finally, a strategy is introduced to merge relevant adjacent clusters and further localize the points where shock waves interact. The Bézier curve fitting technique is then employed to obtain the high-quality shock-lines. Several numerical cases demonstrate that this method achieves high localization accuracy for shock-lines while being minimally affected by grid type and scale variations. Moreover, it enables clear and effective identification of the shock interaction patterns in both steady and unsteady flows, providing an effective visualization means for analyzing the motion and evolution of shock wave configurations.

High-pressure capturing wing (HCW) aerodynamic configuration demonstrates favorable aerodynamic performance under hypersonic conditions, and its novel additional lifting wing (also known as HCW) has the potential to enhance lift characteristics under subsonic conditions. Therefore, this configuration presents a promising option for wide-speed-range vehicles. However, the stability characteristics of this novel configuration under subsonic conditions have not yet been investigated. In this paper, the effects of wing dihedral angles on the subsonic aerodynamic characteristics of a parametric conceptual HCW configuration with two lifting wings were investigated. Specifically, the design variables for this study were the dihedral angles of the upper HCW and the lower delta wing. To obtain the distributions of various aerodynamic parameters over the design space, a combination of the uniform experimental design method, computational fluid dynamics numerical simulation techniques, and kriging surrogate model algorithm was employed. The findings suggest that wing dihedral angles have a greater impact on the lift-drag ratio (L/D) at low angles of attack compared to high angles of attack. L/D can be enhanced by incorporating a positive dihedral angle in HCW, and as the delta wing's negative dihedral angle rises, L/D tends to increase earlier and decrease later at low angles of attack. Furthermore, for the longitudinal, lateral, and directional stability characteristics of this configuration, the positive dihedral angles of the delta wing offer greater overall advantages than negative ones in improving them, and the positive dihedral angles of HCW yield more significant enhancements in stability compared to negative ones.

Shock waves and shock-shock interaction are typical phenomena in supersonic or hypersonic flows that have significant impacts on aerodynamic performance. To obtain a comprehensive understanding of the mechanism of shock wave interaction, shock wave detection (SWD) methods are required. However, it is often challenging for most current SWD methods to identify the relationship between shock waves (also known as shock topology). To address this issue, this paper proposes a novel three-dimensional shock topology detection method based on the tomographic reconstruction strategy. This method involves extracting parallel slices from the flow field, then utilizing a two-dimensional shock topology recognition algorithm to obtain shock lines. Shock bands are obtained by connecting shock lines for every two adjacent slices, and shock surfaces are generated by assembling shock bands. Interaction lines are also formed by connecting interaction points. The detected shock wave is a structure composed of "point-line-band-surface", and the topology relationship with other shock waves is obvious. Numerical results show that the shock waves detected by the proposed method can be categorized into families. Moreover, the shock surfaces generated by this method are free of gaps, holes, and un-physical fragments, which is an improvement over existing SWD methods.

High-pressure capturing wing (HCW) configuration is a potential hypersonic aerodynamic configuration that can simultaneously have good lift-drag characteristics with a large volumetric ratio. The effects of wing dihedral angle on the hypersonic aerodynamic characteristics of a conceptual HCW configuration with two lifting wings are investigated in this paper. Specifically, the dihedral angles of two wings, the upper HCW and the conventional delta wing at the bottom of the body, were regarded as the design variables with a given space. Furthermore, the uniform experimental design method, computational fluid dynamics simulation techniques, and kriging surrogate model algorithm were successively utilized to establish the distributions of aerodynamic parameters over the design space. The results indicate that the lift, drag, and lift-to-drag ratio (L/D) have the similar variation trends as the dihedral angles of the two wings change, and are more sensitive to the positive dihedral angle of the delta wing. When the angle of attack is small, the increasing positive dihedral angle significantly reduces L/D, but as the negative dihedral angle increases, L/D will first increase slightly and then decrease slowly. In particular, when the angle of attack is large, the wing dihedral angles have less influence on L/D. For the longitudinal stability, it is mainly affected by the dihedral angles of the delta wing, and the positive dihedral angle can slightly weaken it, while the negative one hardly changes it. The directional stability can be enhanced by the wing dihedral angles, especially the negative angle. The positive dihedral angle can improve the lateral stability, while the negative weaken it. However, when the angle of attack is large, the large positive dihedral angle of the delta wing may lead to a decrease in the lateral stability.

Based on the conceptual configuration combining the conical-cone airframe and flat plate capturing wing configuration, a series of different shapes were obtained by changing the side expansion angle of the trailing edge.The configuration was solved by computational fluid dynamics under the condition of typical subsonic flow（Ma=0. 5）.The results show that, at the angle of attack of 0°, the spanwise section of the body tail widened, the expansion effect of the channel between the body and the capturing wing on the incoming flow was weakened, and the reverse pressure gradient on the upper surface of the body platform was reduced, which could effectively inhibit the flow separation phenomenon in the channel. In addition, the whole vehicle lift coefficient rose, and the drag coefficient first fell and then grew. With the increase of the angle of attack, the pressure on the upper surface of the round platform of the body climbed, while the range of the separation zone gradually declined until it disappeared. Furthermore, the spanwise widening of the tail section of the body could accelerate the disappearance of the separation zone. When the angle of attack increased further, the transverse flow appeared on the leeward side of the body, while the spanwise widening of the body tail section could delay the development of the transverse flow. As the angle of attack increased, the calculation results also show that the lift and drag of the whole machine were mainly attributed to the capturing wing. The aerodynamic force contributed by the body was not sensitive to the change of the angle of attack. The spanwise widening of the body tail section had little effect on the aerodynamic center of the whole vehicle. The geometry change of the lower surface of the body had no significant effect on the flow characteristics in the channel between the body and the capturing wing and the aerodynamic characteristics of the capturing wing.

This paper proposes a common-weights weighted essentially non-oscillatory (Co-WENO) scheme for solving the Euler equations of gas dynamics. Different from the usual component-wise weighting methods, common-weights means that, on one global stencil, a set of weights is commonly shared by the split flux vector of Euler equations in one spatial dimension. The common-weights WENO scheme has two significant advantages. First, since only one set of weights is calculated and used for the split flux vector, the method has an improved computational efficiency. Second, for a stencil (or each cell on the stencil), the Co-WENO scheme keeps the same contribution on each component numerical flux in a hyperbolic system of equations. How to calculate the weights is one of the vital issues in developing this kind of Co-WENO schemes. In this paper, based on the flux vector split method, the product of density, pressure, and the split flux of energy equation(Γ±=ρpfE±) is proposed to calculate the common weights. This is based on the following considerations: (1) the density jumps at shocks and contact discontinuities; (2) the split energy flux contains the term of the third power of the velocity (for example, u3) and makes the resulting scheme has upwind characteristic; (3) the pressure always jumps at shocks, and it can help improve the stability in high speed flows, in which the kinetic energy is much larger than the internal energy. Numerical experiments also show that the proposed common-weights WENO scheme has good robustness and low numerical dissipation, and it can help suppress phase errors.

The aerodynamic configuration design of wide-range hypersonic vehicle has become one of the research hotspots. The new aerodynamic configuration of high-pressure capturing wing (HCW) can meet the requirements of high volume ra-tio, high lift and high lift-to-drag ratio at the same time. The preliminary research of this configuration is mainly aimed at the hypersonic state. Based on this background, this paper took the wide-range hypersonic vehicle as the main target and developed a new bi-wing configuration according to the basic design principle of HCW. The research results show that the lift coefficient of the vehicle can be increased by about 16.6% when the capturing wing is added in the subsonic range. Furthermore, the capturing wing can suppress the jump of the aerodynamic focus of the vehicle in the transonic range, and the vehicle is statically stable in the wide-speed range.

The waverider is widely used in hypersonic vehicles with its high aerodynamic performance, but due to the serious aerothermal environment, its sharp leading edge should be blunted. Circular blunt is one of the commonly used aerothermal characteristic protection methods. Circular blunt with larger diameter can reduce peak heat flux, but at the same time, it will lead to larger drag. The existing research shows that under the same blunt diameter in two-dimensions, the non-uniform blunt can reduce the peak heat flux by 20%, and the difference of drag is small. In this paper, the non-uniform blunt profile is applied to the three-dimensional waverider, and the influence of the non-uniform blunt profile on the aerothermal characteristic performance and aerodynamic performance of the waverider is studied, and the results are compared with those of circular blunt. The numerical simulation is used to compare and analyze the waverider under different angles of attack, flight altitudes, and Mach number. The results show that the peak heat flux of the waverider with non-uniform blunt reduces by about 17% compared with that with circular blunt under a small angle of attack range, Mach 2-10, and a flight altitude of 15–35 km. Meanwhile, when the blunt height/diameter is 20 mm, the aerodynamic performance difference between the two different blunt profiles does not exceed 3% within a 15 degrees angle of attack, Mach 2-10, and flight altitude of 15–35 km. The non-uniform blunt profile can be applied to the design of the three-dimensional waverider.