Search

As for blunt-body hypersonic vehicles, the extreme drag and aero-heating pose great challenges to the conceptual design. Different from the conventional configurations, a blunt body slotted with a channel is proposed in this paper. The channel is further combined with a jet on the shoulder to reduce drag and aero-heating. Simultaneously, based on the position of the separation point and vortex edge, a modified effective body approach is developed to accurately investigate the drag reduction mechanism. Specifically, the more slender the modified effective body is, the more obvious the drag reduction is. The simulation results show that the spike-channel-jet can push the shear layer on the spike away from the wall. In this way, the position of shear reattachment point moves downstream, which can enlarge the recirculation zone and weaken the shock-shock interaction. Additionally, the reattachment shock is pushed far away from the blunt body, which is also weakened. For drag, net power is introduced to evaluate the practical value of spike-channel-jet for drag reduction. Compared with a simple spike (spike length L / D = 1) of the same size, the spike-channel-jet can reduce the total drag by 24.68 %, the peak pressure coefficient of the blunt body by 38.54 %, the peak temperature of the blunt body by 65.75 %, and the peak temperature of the spike by 10.63 %. The maximum net power can reach 13.30 kW. Considering the fluid-thermal interaction, the findings reveal the influence of the spike length and jet mass flow rate on aerodynamic performance, by modified effective body further analyze the drag reduction mechanism with the jet, and validate the different reasons for the peak temperature and pressure of the blunt body. • Drag and aeroheating reduction mechanism of the spike-channel-jet concept under hypersonic conditions is investigated by the numerical simulation. • The influence factors are analyzed and compared comprehensively. • Combined spike-channel-jet concept is a promising drag and aeroheating reduction technique. [ABSTRACT FROM AUTHOR]

The simulation of thermal–fluid–solid coupling in turbines is critically important for design optimization. Historically, most research on thermal–fluid–solid coupling has been conducted in three-dimensional, often with computational speeds that do not meet practical expectations. This study proposes a one-dimensional performance prediction and multi-objective optimization design methodology for turbines, integrating aerothermodynamics and structural strength, to facilitate rapid multidisciplinary coupling design optimization at a low-dimensional level. Initially, a multidisciplinary coupled turbine performance prediction model is established, incorporating the combined effects of turbine aerothermodynamics and structural mechanics. This model links the thermodynamics of the blade passage with the stress and strain of the blade. The predictive accuracy of this model is validated against experimental data from a four-stage axial flow turbine, showing a maximum error of 1.56% for the total temperature ratio and 1.69% for the total expansion ratio. Subsequently, using blade parameters, degree of reaction, load coefficient, and flow coefficient as optimization variables and targeting the turbine's overall isentropic efficiency and power as optimization objectives, a rapid Non-dominated Sorting Genetic Algorithm II and the Technique for Order Preference by Similarity to an Ideal Solution are employed to optimize the last stage of the four-stage axial flow turbine. The optimized turbine demonstrates an increase in overall isentropic efficiency by 1.333% and an increase in overall power by 3.329%, while satisfying structural strength requirements. The novelty of this study lies in its rapid optimization design and performance prediction method for the coupled aerothermodynamics and structural mechanics at a one-dimensional level. [ABSTRACT FROM AUTHOR]

The circumferential non-uniform tip clearance (CNTC) due to casing out-ofroundness adversely affects the turbine aerodynamic performance due to machining and assembly errors, thermal deformation, and improper active clearance control (ACC), etc. Moreover, the asymmetric computational domain caused by casing out-of-roundness presents difficulties for conventional numerical techniques that consider rotational periodicity. Since previous traditional methods using split computational domains have the disadvantages of high interpolation error and high time cost, an efficient fast-moving mesh (FMM) method based on an algebraic approach is proposed in this paper. This method is first validated by using a single-stage turbine with elliptical casing. The results show that the FMM has the advantages of high accuracy, high efficiency, and easy operation, which helps to solve the CNTC problem quickly in scientific research or engineering applications. Then, the effects of CNTC induced by the elliptical casing on the flow field and aerodynamic performance are investigated by using an in-house code that integrates the FMM method. Finally, the effect of stator row interference on the aerodynamic performance in the turbine stage with an elliptical casing is demonstrated. The results show that different types of elliptical casings have a significant effect on the aerodynamic performance. However, the variation law is not consistent (decreasing by 0.538% or increasing by 0.212%). Importantly, the novel finding of this paper is that this discrepancy is jointly determined by the interaction of multiple secondary flows (passage vortex, scraping vortex, etc.) at different spans, not just related to the variation of the tip leakage vortex (TLV) with tip size. Furthermore, this study is the first to indicate that the stator row interference can mitigate the extent of performance degradation due to elliptical casings by suppressing the development of secondary flows. These results may provide theoretical support for blade tip gap design and can also serve as a reasonable reference for the effective application of ACC in engineering. Finally, low-order harmonic components with high amplitudes are also innovatively found in the rotor row with a CNTC. These components may cause low-engine-order (LEO) resonances that endanger the safe operation of engines. [ABSTRACT FROM AUTHOR]

Procedures of the Direct Simulation Monte Carlo method for weakly ionized flows around reentry vehicles are presented. For ionization and recombination reactions, expressions for the model dependence of the reaction probability on velocities and energies of reagents are derived. An algorithm of dissociative recombination is presented, whose computational efficiency is reached by bypassing the modeling of the interaction of electrons and heavy particles. An approach to the construction of a weighting scheme for elastic collisions and chemical reactions is described, which significantly increases the computational efficiency of the algorithm. An example of using these numerical models and procedures for studying a weakly ionized flow around a reentry capsule under typical reentry co-nditions is given. The computation results are compared to plasma parameters measured in flight experiments. [ABSTRACT FROM AUTHOR]

This article presents a study of aerothermal performance of tubes with regularly-spaced multi-channel twisted tape elements (RS-MTT) installed. This research aimed to find the proper design RS-MTTs that induce swirl flow which potentially improves fluid mixing between the core fluid and the fluid near the tube wall, thereby accelerating the heat transfer rate. Additionally, the effects of Reynolds numbers and free-spacing ratios (s/y) on heat transfer, friction loss, and thermal performance behaviors were examined. The RS-MTTs having different free-spacing ratios (s/y) of 0.0, 0.25, 0.5, 0.75, and 1.0 were tested. Air was utilized as the testing fluid in experiments with Reynolds numbers (Re) spanning from 6000 to 20,000. The utilization of RS-MTTs with s/y = 0.0, 0.25, 0.5, 0.75, and 1.0 augmented heat transfer rates up to 1.74, 1.80, 1.85, 1.90, and 2.15 times given by the plain tube alone while the friction factors increased by 4.22, 4.61, 3.87, and 4.05 times, respectively. At the lowest Reynolds number of 6000, the thermal enhancement factors of the tube containing the RS-MTTs with s/y = 0.0, 0.25, 0.5, 0.75, and 1.0 reached the maximum values of 1.42, 1.35, 1.31, 1.27, and 1.23, respectively. Among the RS-MTTs tested, the RS-MTT with s/y = 0.0 showed the best thermal enhancement factor of 4.56%, corresponding to the heat transfer augmented of 11.52% with a friction penalty of 8.71%. [ABSTRACT FROM AUTHOR]

Advanced aerospace structures are subjected to extremely harsh working environment, including of mechanical, acoustic, thermal, and aerodynamic loads. These combined loads can make the structures exhibit complex response, which has already attracted increasing concern in the design of advanced aerospace structures. To tackle the problem, a finite element model (FEM) by discretizing the theoretical differential equation is established to calculate the dynamic response of thin-walled structures under combined thermal and acoustic loads. The numerical analysis indicates that three types of nonlinear responses of the thin plate under combined thermal-acoustic loads are obtained, and the mechanism of snap-through is revealed through two ways: thermal buckling analysis and thermal modal analysis. which can be used to explain the nonlinear response characteristics. [ABSTRACT FROM AUTHOR]

This paper presents a new force measurement technique to investigate the effect of aerothermodynamic phenomena, particularly aerodynamic heating, on aerodynamic forces. This technique is successfully applied within a hypersonic combined test facility that integrates arc-jet and shock tunnel. A preheated model transport system is developed to axially immobilize the model during preheating, safe transport to the shock tunnel, and ensure quasi-free-body axial motion during force measurement tests. A comparative evaluation revealed a significant 11.2 % increase in drag coefficient for test models preheated in the arc-jet tunnel, followed by an 8.56 % decrease when cooled to room temperature. By closely comparing drag coefficients under three conditions (cold, hot, and cooled), this study analyzes the distinct effects of ablation-induced shape change and surface temperature on drag coefficients, respectively. This technique allows for a more realistic simulation of hypersonic flight conditions within ground test facilities. [ABSTRACT FROM AUTHOR]

This paper investigates the role of magnetohydrodynamics (MHD) on the aerothermodynamics (ATD) of a representative entry vehicle while flying into the Martian atmosphere. By strategically placing a flight-ready superconducting magnet at varied positions in the Schiaparelli reentry capsule of the ExoMars mission, we discern its impact on essential flow properties. The primary consequence of MHD during atmospheric entry is the generation of the Lorentz force, which increases the shock standoff distance resulting in a reduction of the heat flux on the spacecraft by pushing high-energy plasma particles away. Through different magnet configurations, three distinct cases are formed to comprehensively understand the effects and implications of each setup. The study is performed using the COOLFluiD MHD for EnTries, an in-house ATD solver. For case 1, the magnet's placement behind the ExoMars forebody at the stagnation point reduces the heat flux. In case 2, the magnet's relocation to the shoulder region explores its potential to mitigate communication blackouts by influencing the wake region's flow. However, this positioning also induces shock bending, leading to variations in post-shock species mass fractions and heat flux spikes in the post-shock region. Case 3, involving an additional magnet where the shock bends in case 1, showcases a consistent increase in shock standoff distance across the forebody, providing a longer relaxation zone for species equilibration. Our findings highlight that while the strength of the applied magnetic field is crucial, the magnet's size is equally pivotal in determining ATD behavior. Case 3 emerges as the most promising configuration, consistently reducing heat flux across the forebody and maintaining it in the afterbody. This study underscores the potential of multi-magnet configurations as next-generation MHD heat shields for Martian atmospheric entry, emphasizing the criticality of magnet placement and configuration in enabling future MHD-enhanced deep space exploration missions. [ABSTRACT FROM AUTHOR]

The real-time monitoring of heat fluxes on the surface of a flight vehicle is vital for safe reentry and efficient attitude control. However, conducting measurements during flight is a challenge. In this study, a real-time heat flux estimation method using inverse heat conduction was proposed, and internally mounted sensors were used for measurements. First, the time-varying samples of heat flux on the surface and temperature on the inner wall were generated along a variety of reentry trajectories. Second, a sensor selection algorithm based on feature importance was applied to select optimal sensor mounting locations. Finally, a prediction model was built using the random forest algorithm to estimate surface heat flux with measured temperatures from the mounted sensors. The proposed method was employed to predict the real-time heat flux of a reentry capsule during return flight. The results show that the proposed method can predict heat flux in real time with a prediction error of less than 0.2%. Further, the sensor selection algorithm enhanced prediction efficiency by reducing the number of necessary sensors. [ABSTRACT FROM AUTHOR]

An analysis of main features of unsteady heat flux measuring in shock tubes experiments with a characteristic process time of ~1 μs–1 ms using sensors based on anisotropic bismuth thermoelements is made. The heat flux behind the reflected shock wave and at the blunt body stagnation point has been measured. Testing of the heat flux calculation method was carried out. The difference between the experimental data and the theoretical value of the heat flux does not exceed 50%. The possible reasons influencing the magnitude of measurement uncertainty are analyzed. The experiments performed have shown the applicability of a sensor based on anisotropic thermoelements and the method for calculating the heat flux from its electrical signal for typical conditions of experiments on shock tubes. [ABSTRACT FROM AUTHOR]

A magnetic field might interact with its surroundings, influencing chemical and physical processes in materials processing, heat exchangers, and other scientific study. Therefore, a computational study of non‐Newtonian (Casson) free convective MHD unsteady fluid flow has been highlighted in this article with mass and heat transit property through a vertical infinite porous plate. A sinusoidal boundary conditions as well as chemical reaction and thermal radiation have been considered. Using a collection of nondimensional variables, the flow related equations are also turned into nondimensional form. The EFDM algorithm is employed in order to arrive at a numerical solution via Compaq Visual Fortran. The reliability of the numerical solution has been confirmed using stability testing and convergence analysis. The whole system is convergent when the values of Prandtl number and Lewis number are greater than or equals to 0.075 and 0.025, respectively. A visual depiction of the impact of the pertinent factors on dimensionless velocity, temperature, and concentration profiles are displayed through graphical representation as well as with tabular representation. It has been inspected that when the magnetic component is regarded, it greatly affects the heat transfer factors of Casson nanofluid and the heat also rises when Eckert number, heat source and radiation parameter accelerate. It is also found that the Sherwood number is increased as the impact of chemical reaction parameter and the Lewis number, also the skin friction is decreased as the influence of porosity term got accelerated. The comparison of the current findings with the data that were previously published serves as the final stage in validating the present study. [ABSTRACT FROM AUTHOR]

By slotting an inlet at the head of the aerodisk, a channel through the spike, and the lateral jet outlet at the middle of the spike, the blunt body with spike-aerodisk can further reduce its aeroheating and drag. In this paper, the numerical simulation method was used to analyze the flow field structure of this spike-aerodisk-channel, the effects of the spike length, the aerodisk diameter, and the channel diameter on the drag and aeroheating performance of this spike-aerodisk-channel configuration were studied. In general, the lateral jet introduced by the channel from the stagnation region of the aerodisk leads to the shear layer separation in advance, pushes the separation shock away from the spike, enlarges the recirculation zone near the blunt body, avoids the interaction between the separation shock and the reattachment shock, increases the distance from the reattachment shock to the surface of the blunt body. The reattachment shock strength of the blunt body is weakened. With increased spike length and channel diameter, drag and aeroheating reduction are improved. With the increase of the diameter of the aerodisk, the aeroheating performance is improved, but the drag reduction performance continues to decline. Compared with the plain spike-aerodisk, the spike-aerodisk-channel is generally conducive to further drag and aeroheating reduction. In the range of research parameters, considering the effect of drag and aeroheating reduction and the structural feasibility, the spike-aerodisk-channel has the best drag and aeroheating reduction performance when the spike length is twice the blunt-body diameter, the ratio of the aerodisk diameter to the blunt-body diameter is 0.3, and the ratio of the channel diameter to the blunt-body diameter is 0.075. Compared with the plain spike-aerodisk of the same size, the aeroheating reduction efficiency is increased by 37.9%, and the drag reduction efficiency is increased by 16.0%. • Drag and aeroheating reduction mechanism of the spike-aerodisk and channel concept under hypersonic conditions was investigated by the numerical simulation. • Three influence factors were analyzed and compared comprehensively. • Combined spike-aerodisk and channel concept is a promising drag and aeroheating reduction technique. [ABSTRACT FROM AUTHOR]

The efficiency of various industrial processes is contingent upon the heat transfer enhancement and fluid flow dynamics within horizontal annular heat exchangers. This study, therefore, investigates these factors, utilizing a heater with the capacity to vary between 400 W and 600 W, connected to an alternating current (AC) voltage regulator. Cold water, with flow rates ranging from 2.5 to 5 gallons per minute (GPM), circulates in a closed-loop system. The central objectives of this research are to elucidate the heat transfer characteristics, including heat transfer coefficient, Nusselt number, friction factor, and flow regime, and to derive a new empirical correlation for the Nusselt number (Nu) and the friction factor (f) within a horizontal annular heat exchanger incorporating a central tubular heater. The experimental evidence demonstrates an increase in the heat transfer coefficient (h) and the Nusselt number (Nu) with the rise in Reynolds number (Re). The peak heat transfer coefficient was observed at a heat rate of 450 W, commencing at 1893 W/m2·K when Re=4813 and reaching a maximum of 3886 W/m2·K when Re=10127. Additionally, the mean Nusselt number (Nu) was found to range between 89 at Re=4813 and 183 at Re=10127. Conversely, the friction factor (f) decreased as the Reynolds number (Re) increased. The transition from a laminar to a transitional flow regime was observed at Re=8000. The empirical correlations derived from this study were Nu = 0.145Re0.671Pr0.4 and f = 1.4219Re-0.277 with mean absolute percentage errors (MAPEs) of 10.64% and 25.70%, respectively. These findings have the potential to significantly enhance the performance of horizontal annular heat exchangers in various industrial applications. [ABSTRACT FROM AUTHOR]

A large amount of heat flux from aerodynamic heating acts on reusable spacecraft; thus, an accurate heat flux prediction around spacecraft reentry is essential for developing a high-performance reusable spacecraft. Although the approximate convective-heating equations can calculate the heat flux with high efficiency and sufficient fidelity, the heat flux should be evaluated over a thousand times for the entire trajectory in multidisciplinary analyses. For these reasons, it is necessary to develop an efficient method for calculating the heat flux for multidisciplinary analysis. In this paper, an efficient method for heat flux calculation that is adoptable by multidisciplinary analyses for hypersonic vehicles, such as spacecraft, is developed. Approximate convective-heating equations were adopted to relieve the computational cost of estimating the heat flux, and an adaptive time step method for heat flux calculations was developed to reduce the number of heat flux calculations required across the entire flight trajectory. A dynamic factor was introduced that adjusts the time step between each instance of the heat flux calculation. Since the time step using this factor could increase under low heat flux conditions, the number of heat flux calculations decreases by approximately one-tenth with over 90% accuracy. Therefore, the efficiency was improved with high accuracy using the adaptively-determined time step according to this dynamic factor. [ABSTRACT FROM AUTHOR]

Photon trapping is believed to be an important mechanism in super-Eddington accretion, which greatly reduces the radiative efficiency as photons are swallowed by the central black hole before they can escape from the accretion flow. This effect is interpreted as the radial advection of energy in one-dimensional height-integrated models, such as the slim-disk model. However, when multidimensional effects are considered, the conventional understanding may no longer hold. In this paper, we study the advective energy transport in super-Eddington accretion based on a new two-dimensional inflow–outflow solution with radial self-similarity, in which the advective factor is calculated self-consistently by incorporating the calculation of radiative flux instead of being set as an input parameter. We found that radial advection is actually a heating mechanism in the inflow due to compression, and the energy balance in the inflow is maintained by cooling via radiation and vertical (θ -direction) advection, which transports entropy upward to be radiated closer to the surface or carried away by the outflow. As a result, fewer photons are advected inward, and more photons are released from the surface, so that the mean advective factor is smaller and the emergent flux is higher than the fluxes predicted by the slim-disk model. The radiative efficiency of super-Eddington accretion thus should be higher than that of the slim-disk model, which agrees with the results of some recent numerical simulations. [ABSTRACT FROM AUTHOR]

Designing spacecraft involves a careful equilibrium to avoid overengineering or underdesigning, which underscores the importance of employing thermal uncertainty analysis. A key part of this analysis is modeling thermal conditions, but this is often a computationally heavy process. This is largely because ray-tracing calculations require determining the external heat flux of solar radiation across different operating conditions. Ray emission varies across conditions, which can lead to inefficient resource use in uncertainty calculations. Our study aims to address this by introducing a new approach to calculating the external heat flux of solar radiation that is better suited for uncertainty analysis than previous approaches. Our formula only requires ray tracing to be performed for one condition rather than for every condition. The other conditions are handled by simple matrix budgeting, negating the need for complicated ray tracing. In the aforementioned analytical procedure, certain matrices demonstrate sparsity properties. By exploiting this characteristic, optimization computations can be executed by utilizing sparse matrix algorithms. We tested this new formula, which we call the external heat flux expansion (EHFE) formula, on a specific spacecraft and compared the results with those obtained using the traditional method. Our findings suggest that the EHFE formula is ideal for calculating uncertainty. It significantly improves computational efficiency while maintaining accuracy. The formula is also user-adjustable, allowing the accuracy of uncertainty calculation results of the external heat flux of solar radiation to be fine-tuned by changing the value of the cutoff factor. This work establishes an essential theoretical framework pivotal to addressing inherent uncertainties in the thermal design of upcoming deep-space exploration spacecraft, solar observatory satellites, and space solar power stations. [ABSTRACT FROM AUTHOR]

A high-temperature plasma sheath is generated on the surface of the re-entry vehicle through the conversion of kinetic energy to thermal and chemical energy across a strong shock wave at the hypersonic speed. Such a condition results in the forming of a blackout which significantly affects the communication components. To analyze the re-entry vehicle at the hypersonic speed, an unconditionally stable system incorporated factorization-splitting (SIFS) algorithm is proposed in finite-difference time-domain (FDTD) grids. The proposed algorithm shows advantages in the entire performance by simplifying the update implementation in multi-scale problems. The plasma sheath is analyzed based on the modified auxiliary difference equation (ADE) method according to the integer time step scheme in the unconditionally stable algorithm. Higher order perfectly matched layer (PML) formulation is modified to simulate open region problems. Numerical examples are carried out to demonstrate the performance of the algorithm from the aspects of target characteristics and antenna model. From resultants, it can be concluded that the proposed algorithm shows considerable accuracy, efficiency, and absorption during the simulation. Meanwhile, plasma sheath significantly affects the communication and detection of the re-entry vehicle. [ABSTRACT FROM AUTHOR]

Thermal uncertainty analysis of spacecraft is an important method to avoid overdesign and underdesign problems. In the context of uncertainty analysis, thermal models representing multiple operating conditions must be invoked repeatedly, leading to substantial computational costs. The ray tracing calculation of Earth infrared and albedo radiation heat flux is an important reason for the slow calculation speed. As the rays emitted during external heat flux calculations under different operating conditions are independent and unconnected, the rays produced across various conditions are effectively wasted. In this study, the external heat flow equation is thoroughly expanded and the derived factors are clustered and analyzed to develop a novel formula for calculating external heat flow. When this formula is employed to compute the uncertain external heat flux, only one condition necessitates ray tracing, while the remaining conditions utilize simple matrix operations in place of complex ray tracing. Within the aforementioned procedure, certain matrices demonstrate sparse characteristics. The optimization calculations for these matrices can, therefore, benefit from the application of sparse matrix optimization algorithms. Using a spacecraft as an example, the uncertain external heat flux calculation outcomes of the new and traditional formulas are compared and assessed. The findings reveal that the new formula is highly suitable for estimating uncertain Earth radiation heat flow, with a marked improvement in efficiency. The accuracy is essentially equivalent to that of the traditional formula and the calculation precision can be dynamically adjusted to meet user requirements. The methodology can be further generalized to assess the uncertainties associated with radiative external heat fluxes for other celestial bodies within the solar system. This offers a valuable theoretical framework for addressing the uncertainties in the thermal design of deep space exploration vehicles. [ABSTRACT FROM AUTHOR]

An analytical trajectory prediction method is put forward for skip re-entry of lifting vehicles. First, the 3-D analytical solutions of skip re-entry trajectory are developed, which are crucial to realize the analytical trajectory prediction. To facilitate the derivation, an auxiliary geocentric inertial frame is established to deal with the strong nonlinear problem caused by the Earth's rotation. Thereafter, the analytical solutions of the downrange, crossrange, and velocity are approximated by the outer solutions of a two-time scale singular perturbation. The analytical solutions of the altitude and flight-path angle are developed using regular perturbation techniques which can decompose the strongly coupled and nonlinear system into a sequence of linear problems. Second, since the analytical solutions are with respect to the lift-to-drag ratio and bank angle, a method based on the EKF technique is proposed to estimate the lift-to-drag ratio and bank angle of the re-entry vehicle. Third, the analytical trajectory prediction schemes for the lifting re-entry vehicle with multiple bank reversals are put forward. The simulations show that the proposed analytical solutions are in good agreement with the numerical solutions, but the computation time of the analytical solutions is about two orders of magnitude less than that of the numerical solutions. In addition, the superior performance of the proposed analytical trajectory prediction method is well demonstrated by the simulation results. [ABSTRACT FROM AUTHOR]

The thermal comfort and air-conditioning energy consumption of vehicles or trains are dependent on the thermal resistance of its envelopes, which could be enhanced by improving the radiation characteristics of the narrow cavities scattered in their envelopes. However, the study for a feasible method and its effectiveness has been given little attention. This paper introduces a method that involves pasting aluminum foil on the inner wall to change the radiation characteristics of the narrow cavity and analyzed its effects on narrow-cavity heat transfer by experimental and numerical methods. The results indicate that the radiation effect on heat transfer in a narrow cavity made of conventional material is dominant, with a rate larger than 75%, and that pasting aluminum foil is an effective and feasible method of weakening the radiation rate in narrow-cavity heat transfer, decreasing it to less than 10%. This paper will provide a reference to improve the insulation characteristics of vehicle enclosure. [ABSTRACT FROM AUTHOR]

We developed and tested two carbon–phenolic-based ablators for future Korean spacecraft heat shield applications. The ablators are developed with two layers: an outer recession layer, fabricated from carbon–phenolic material, and an inner insulating layer, fabricated either from cork or silica–phenolic material. The ablator specimens were tested in a 0.4 MW supersonic arc-jet plasma wind tunnel at heat flux conditions ranging from 6.25 MW/m2 to 9.4 MW/m2, with either specimen being stationary or transient. Stationary tests were conducted for 50 s each as a preliminary investigation, and the transient tests were conducted for ~110 s each to stimulate a spacecraft's atmospheric re-entry heat flux trajectory. During the tests, each specimen's internal temperatures were measured at three locations: 25 mm, 35 mm, and 45 mm from the specimen stagnation point. During the stationary tests, a two-color pyrometer was used to measure specimen stagnation-point temperatures. During the preliminary stationary tests, the silica–phenolic-insulated specimen's reaction was normal compared to the cork-insulated specimen; hence, only the silica–phenolic-insulated specimens were further subjected to the transient tests. During the transient tests, the silica–phenolic-insulated specimens were stable, and the internal temperatures were lower than 450 K (~180 °C), achieving the main objective of this study. [ABSTRACT FROM AUTHOR]

This article presents an indirect neural-based finite-time integral sliding mode control algorithm for the reference trajectory tracking guidance of Mars entry vehicle under uncertainties. The proposed controller is developed as a combination of finite-time integral sliding mode controller and indirect neural identification. The finite-time integral sliding mode controller is designed by constructing a new type of finite-time integral sliding mode surface to prevent the singularity problem. Moreover, the neural network (NN) is combined with the finite-time integral sliding mode controller to identify the lumped uncertainty and attenuate the chattering phenomenon. Particularly, the concept of indirect neural identification is adopted and only a single adaptive parameter is required to be learned online. In this way, the proposed controller is not only strongly robust against aerodynamic and density uncertainties, but also computationally simple for onboard implementations. Stability argument indicates that the proposed controller can ensure the radial distance tracking error and its time differentiation regulate to the small residual sets around zero in finite time. Lastly, the effectiveness and excellent guidance performance of the proposed control algorithm are demonstrated through simulation studies on a Mars Science Laboratory-type (MSL-type) entry vehicle. [ABSTRACT FROM AUTHOR]

The bang-bang control type of bank angle solution is discussed in many studies that address the Mars entry guidance problem for the maximum terminal altitude. Based on existing evidence, this paper proposes a novel entry guidance method that can generate a suboptimal entry trajectory governed by a predetermined bank angle profile designed to reflect bank rate constraint and crossrange control. This converts the original entry optimization problem into a parameter optimization problem with four time variables that need to be determined. Since the minimum terminal velocity problem has the similar necessary condition of optimality as the maximum terminal altitude problem, the predetermined bank angle profile can also be applied to the minimum terminal velocity problem. The entry phases of Mars robotic and human missions are adopted as examples to verify the performance of the proposed guidance method. Simulation results demonstrate that the proposed guidance method can generate a suboptimal entry trajectory at which the final state is comparable to that of a nonlinear programming solver. Importantly, the proposed guidance method demonstrates high computational efficiency so that it can be implemented as an onboard guidance method. [ABSTRACT FROM AUTHOR]

The exergy analysis has already proved to be of interest for aerospace applications, with the possibility to analyze new types of configurations. It is of particular interest when a pertinent thrust/drag breakdown is not possible and when significant thermal exchanges take place in the control volume. In addition, this method allows to link the different contributions of the exergy terms to physical phenomena in the flow, improving the understanding of the sources of losses degrading the performance of the system under study. The applications and results obtained in a fixed frame of reference have thus motivated the improvement and extension of the exergy balance to more complex flows. This paper aims at presenting an extension of the exergy balance to rotating frames of reference for the analysis of rotating components of the propulsion system (e.g. propeller, turbomachine or helicopter rotor blades). The newly developed theoretical development is summarized, and an application to a numerical simulation of a propeller is presented. Results are discussed in terms of both physical interpretation and numerical accuracy. [ABSTRACT FROM AUTHOR]

Purpose: Development or upgradation of airplanes requires many different analyses, e.g. thermal, aerodynamic, structural and safety. Similar studies were performed during configuration change design of commuter category aircraft equipped with pusher turboprop engines. In this paper, thermo-fluid analyses of interactions of the new propulsion system in tractor configuration with selected elements of airplane skin are carried out. This study aims to check the airplane skin material, and its geometry, including the Plexiglas passenger window material degradation, due to hot exhaust gas plume impingement. The impact of change in exhaust stub angle and asymmetric inboard-outboard stubs on the jet thrust at various flight operating conditions like minimum off-route altitude and cruise performance is assessed. Design/methodology/approach: Commercial software-based numerical models were developed. In the first stage, heat and fluid flow analysis was performed over a twin-engine airplane's nacelle, wing and center fuselage with its powerplant mounted in the high wing configuration. Subsequently, numerical simulations of thermal interactions between the hot exhaust gases, which leave the exhaust system close to the nacelle, flaps and the center fuselage, were estimated for various combinations of exhaust stub angles with asymmetry between inboard-outboard stubs at different airplane configurations and operating conditions. Findings: The results of the simulations are used to recommend modifications to the design of the considered airplane in terms of material selection and/or special coatings. The importance and impact of exhaust jet thrust on the overall aircraft performance are investigated. Originality/value: The advanced numerical model for the exhaust jet-airplane skin thermal interaction was developed to estimate the temperature effects on the propeller blades and aircraft fuselage surfaces during different flight operating conditions with multiple combinations of stub orientations. [ABSTRACT FROM AUTHOR]

In addition to the flight safety of re-entry high speed commercial air vehicles their environmental eligibility come into particular important issue due to environmental requirements for commercial subsonic aviation are becoming more stringent. During the re-entry take-off and approach cycles the vehicles are a subject to almost the same type of environmental requirements as the commercial subsonic aviation. The requirements are primarily concern the limitations of community noise and engine exhaust emissions levels, as well as the sonic boom level (for commercial supersonic aviation). Meeting the requirements may significantly complicate the development, certification and further operation of the vehicles. The paper discusses one of the potential ways of increasing the acoustic eligibility of the vehicles during take-off and approach cycles at using the some number of turbofan engines as part of the propulsion system. The way is based on the creation of a special control of turbofan engines thrust/power (so called low noise thrust management) accounting the provision of flight safety restrictions. The assessment of the acoustic impact (community noise) of the vehicle is carried out taking into account the contribution of major noise sources of turbofan engines, i.e. the fan and exhaust jet. The impact of two parameters of turbofan engines such as the bypass ratio and the number of turbofan engines was evaluated. The variation of these parameters was performed under the condition meeting the specified mission requirements in term of the required field (runway) length and the approach speed. The estimations shown that using the created low noise turbofan engine thrust management at the take-off and approach may significantly increase the acoustic eligibility of the vehicles using turbofan engines. • Thrust management may significantly reduce the vehicle noise at take-off and approach. • Low noise thrust management should account the vehicles flight safety restrictions. • Change of turbofan engines number may improves noise and flight safety of vehicle. • Change of bypass ratio/engine number should performed at keeping mission requirements. • Increased glide angle at low noise management requires automatic instrumental system. [ABSTRACT FROM AUTHOR]

The interest of the scientific community in Titan, Saturn's largest moon, is still strong. In fact, after the successful Cassini-Huygens mission, which landed the Huygens capsule on Titan, other missions are planned in the coming years. The aims of this work are to provide the extent of global aerodynamic force, to provide thermal and aerodynamic loads on a capsule entering Titan atmosphere, and to evaluate the effects of chemistry on 1) flowfield parameters, 2) surface catalyticity, 3) temperature and pressure distributions on a heat shield, and 4) global aerodynamic force. To develop this study, the authors used the Huygens capsule and related entry trajectory. The study was carried out in the altitude interval of 295-470 km by means of the direct simulation Monte Carlo codes: DS2V for the solution of two-dimensional/axisymmetric flowfields, and DS3V for the solution of three-dimensional flowfields. The Titan atmosphere is a mixture of nitrogen, methane, and argon. The chemical model is made of 221 reactions, making gas around the capsule a mixture of 18 species. Chemistry strongly influences local aerodynamic quantities, both in the flowfield and on the capsule surface; whereas the chemical effects on the global aerodynamic force are negligible. Our computations also verified the lack of ionization. [ABSTRACT FROM AUTHOR]

Aero-propulsion, in particular applications involving hypersonic flight, benefits from ultra-high temperature ceramics (UHTC) efficiency. Atomistic simulations were carried out to beter understand the thermomechanical characteristics of these materials at the elevated temperatures and pressures of hypersonic flight. To determine the first oxidation of α-SiC (silicon carbide), both the classical molecular dynamics simulation code LAMMPS (Large-scale Atomic/Molecular Massively Parallel Simulator) and the reactive ReaxFF potential were used. The ablation of SiO3, CO, CO2, C2O2, and C2O3 molecules in these simulations provide evidence of the first ablative characteristics of α-SiC as well as the distinctive parabolic reaction kinetics found in very harsh settings. The temperature range selected for these simulations was 26 °C to 2,000 °C. Due to the unique oxidation kinetics present in these extreme environments, further research into the oxidation of UHTC would require the use of newly parameterized reactive potentials. This work assesses the effect of extreme environments and the volatilization of molecules from polycrystalline surfaces as a result of the unique oxidation kinetics that are present in these extreme environments. [ABSTRACT FROM AUTHOR]

The mechanical expansion reentry vehicle has become the focus of deep space exploration because of its good deceleration effect and high stability. However, due to its special aerodynamic shape, its surface heat flux characteristics are different from traditional reentry vehicles. In this paper, the Two-Temperature model is introduced to simulate heat flux distribution. The influence of different structure parameters and flight parameters on the flow field structure and surface heat flux is also analyzed. The research shows that the Two-Temperature model can improve the prediction accuracy and that the heat flux may peak at the both the head and shoulder of the vehicle. Structural parameters RB, RN, and θ have an obvious negative effect on QO. RB, RN, RR, and LZ have a negative correlation with QR. QR drops first and then rises as θ increases and RS decreases. Flight parameters Ma have a positive effect on QO and QR while H is negative; α makes the heat flux distribution asymmetric. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

The OSIRIS-REx spacecraft completed the first part of the primary objective by successfully sampling the surface of asteroid (101955) Bennu and storing the acquired sample in the re-entry capsule. The sampling 'Touch-And-Go' (TAG) maneuver was performed nominally at the primary sampling site, Nightingale, in Bennu's Northern hemisphere. As a consequence of the TAG, material at the sampling site was mobilized and the morphology of the area was altered. This event offered a unique opportunity to investigate, in detail, the subsurface of asteroid Bennu giving access to fine grained and less altered material from Nightingale crater. We performed a detailed study on the infrared spectrum in the Nightingale region to search for modification resulting from the sampling event by analysing different features: slope, the H2O–OH− related absorption band in the |$2.7\ \mu {\rm m}$| region, and other possible features. Our results show that, despite visible alteration of the TAG location detected by cameras, no strong variations are observed in the near-infrared bands and their quantitative evaluation is not possible beyond all the instrumental effects, although some changes may have occurred. We confirm that the infrared spectrum of the sampling site becomes redder with respect to the pre-TAG observations, conceivably due to fine material mobilization and exposure of less altered material, as confirmed by decreased spectral convexity. We identify possible modification of hydrated band at |$2.7\ \mu {\rm m}$| but with some concerns due to data quality. However, our results place new constraints on the nature of Bennu's subsurface material and the sample collected by OSIRIS-REx. [ABSTRACT FROM AUTHOR]

BACKGROUND: High-G acceleration experienced during launch and re-entry of suborbital spaceflights may present challenges for older or medically susceptible participants. A detailed understanding of the associated physiological responses would support the development of an evidence-based medical approach to commercial suborbital spaceflight. METHODS: There were 24 healthy subjects recruited into 'younger' (18-44 yr), 'intermediate' (45-64 yr) and 'older' (65-80 yr) age groups. Cardiovascular and respiratory variables were measured continuously during dynamic combinations of +Gx (chest-to-back) and +Gz (head-to-foot) acceleration that simulated suborbital G profiles for spaceplane and rocket/capsule platforms. Measurements were conducted breathing air and breathing 15% oxygen to simulate a cabin pressure altitude of 8000 ft. RESULTS: Suborbital G profiles generated highly dynamic changes in heart rate, blood pressure, and cardiac output. G-induced hypoxemia was observed, with minimum arterial oxygen saturation < 80% in a quarter of subjects. Increased age was associated with greater hypoxemia and reduced cardiac output responses but did not have detrimental cardiovascular effects. ECG changes included recurrent G-induced trigeminy in one individual. Respiratory and visual symptoms were common, with 88% of subjects reporting greyout and 29% reporting blackout. There was one episode of G-induced loss of consciousness (G-LOC). DISCUSSION: Suborbital acceleration profiles are generally well tolerated but are not physiologically inconsequential. Marked hemodynamic effects and transient respiratory compromise could interact with predisposing factors to precipitate adverse cardiopulmonary effects in a minority of participants. Medically susceptible individuals may benefit from expanded preflight centrifuge familiarization that includes targeted physiological evaluation in the form of a 'G challenge test'. [ABSTRACT FROM AUTHOR]

The design and optimization of re-entry spacecraft or its subsystems is a multidisciplinary or multiobjective optimization problem by nature. Multidisciplinary design optimization (MDO) focuses on using numerical optimization in designing systems with several subsystems or disciplines that have interactions and independent actions. In the present paper, the system-level optimizer, trajectory, geometry and shape, aerodynamics, and aerothermodynamics differential equations, are converted to algebraic equations using the Radau pseudospectral method (RPM) since a spacecraft is a nonlinear, extensive, and sparse system. The solution to the problem with the help of MDO is reached by iterating all the disciplines together; one can simultaneously enhance the design, decrease the time and cost of the entire design cycle, and minimize the structural mass of a re-entry spacecraft. Considering various methods presented in earlier research works, a combined and innovative all-at-once (AAO), RPM-based MDO method, including the key subsystems in the design process of a re-entry capsule-shape spacecraft with a low lift-to-drag ratio (L/D), is presented. Considering the applicable state and control variables, various constraints, and parameters applied to several geometric shapes of a blunt capsule and using Apollo's aerodynamic and aerothermodynamic coefficients, the optimized dimensions for a re-entry spacecraft are presented. The introduced optimization scheme led to a 17% mass reduction compared to the original mass of the Apollo vehicle. Fast computing and simplified models are used together in this method to analyze a wide range of vehicle shapes and entry types during conceptual design. [ABSTRACT FROM AUTHOR]

In the last decade, the increasing use of NanoSats and CubeSats has made the re-entry capsule an emerging research field needing updates in configuration and technology. In particular, the door to advancements in terms of efficiency and re-usability has been opened by the introduction of inflatable and/or deployable aerodynamic brakes and the use of on-board electronics for active control. Such technologies allow smaller sizes at launch, controlled re-entries, and safe recovery. This paper deals with the design of a guidance and control algorithm for the re-entry of a capsule with a deployable aero-brake. A trajectory optimization model is used both in the mission planning phase to design the reference re-entry path and during the mission to update the trajectory in case of major deviations from the prescribed orbit, thanks to simplifications aimed at reducing the computational burden. Successively, a trajectory tracking controller, based on Nonlinear Model Predictive Control (NMPC), is able to modulate the opening of the aero-brake in order to follow the planned trajectory towards the target. A robustness analysis was carried out, via numerical simulations, to verify the reliability of the proposed controller in the presence of model uncertainties, orbital perturbations, and measurement noise. [ABSTRACT FROM AUTHOR]

The characteristics of aero-thermoelastic coupling are important for the design of the leading edge in hypersonic vehicles. Herein, a fluid–structure interaction analysis is performed to study the leading edge of a hypersonic vehicle using aero-thermoelastic coupling methods. The results show that the maximum heat flux and temperature of the optimized Bézier curve leading edge are reduced to a certain extent, compared with a hemi-cylindrical leading edge, and the lift–to–drag ratios of the two models are close. The Bézier curve leading-edge model can reduce the blunt radius of the leading edge of the hypersonic vehicle and increase the aerodynamic performance without losing thermal performance. [ABSTRACT FROM AUTHOR]

High-speed aerospace applications, such as re-entry vehicles, mostly involve thin-walled structural components with a high strength-to-weight ratio and high-temperature resistant. The present novel work comprises the structural and thermal analysis of re-entry vehicle nose structures made of four functionally graded materials (FGM). Four FGM shell structures made of aluminum/silicon carbide, aluminum/aluminum oxide, Ti-6Al-4V/silicon carbide and Ti-6Al-4V/aluminum oxide have been considered for the re-entry vehicle nose. The effect of various thermal environments, as well as the linear temperature rise from metal-rich to ceramic-rich on critical buckling temperature and natural frequency have been studied. The critical buckling temperature, as well as the natural frequency of the large, thin re-entry vehicle nose structures, decrease with an increase in a uniform thermal environment, as well as linear temperature rise. The effect of shell thickness on buckling and dynamic characteristics of an FGM shell is also studied, suiting the nose of the re-entry vehicle under various linear temperature rises. The critical buckling temperature and natural frequency are quantified for several cases, and it was observed that they are significantly influenced by the shell thickness. Thus, the research intends to determine the thickness required for such thin and large shells to withstand in the re-entry thermal conditions. [ABSTRACT FROM AUTHOR]

The use of passive, active, or hybrid flow control techniques is often investigated to reduce the acoustic signature, wave drag, and aerodynamic heating associated with the supersonic flow regime. This research explores passive and hybrid flow control techniques to achieve an optimal reduction in wave drag and aerodynamic heating on a blunt body using an aerodisk. While passive techniques use one or two aerospikes, hybrid techniques employ opposing jets and aerospikes. Numerical analysis is performed using Reynolds-Averaged Navier–Stokes (RANS) equations to analyze the bodies' flow field. The statistical technique, Design of Experiments (DOE), is combined with Response Surface Method (RSM) to find the optimal configuration for four cases by generating design space. Two cases were considered for the optimization: single aerospike with and without opposing jet and double aerospike with and without opposing jet. Variables used for the design of the aerodisks were spike length and diameter, while the response variables were wave drag and normalized heat flux. The current study has established an optimum relationship between spike length and aerospike diameter located in front of the main blunt body for both single and double aerospikes. The study's results suggest that a double aerodisk configuration is more beneficial for reducing drag and heat flux at supersonic speed than a single aerodisk. By incorporating an opposing jet at a pressure ratio of 0.8 from the frontal aerodisk to the spiked blunt body, it can reduce drag and heat flux by 86% and 95%, respectively. Finally, numerical verification is performed for statistically optimized designs. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

This document presents the results of a coupled fluid-thermal simulation along the trajectory of a hypersonic vehicle for different surface coatings. This kind of simulations will become more and more important in a reliable design of the thicknesses of the Thermal Protection System (TPS) of hypersonic vehicles. The considered vehicle is a suborbital payload carrier launched by a Lorentz Rail Accelerator (LRA). It is non-propelled and reaches an altitude of 115 km, where it leaves its payload. To reach the desired altitude, the payload carrier has to be accelerated at the Rail reaching a Mach number of 6,2 at its outlet. To determine the temperature evolution of the vehicle, the accurate simulation of the first instants of the trajectory is of outmost importance. Therefore, a coupled simulation provides a better estimation of the maximal temperature, necessary to dimension the TPS. [ABSTRACT FROM AUTHOR]

In this work, an innovative lobe-type rear variable area bypass injector is proposed, which can take into account both mixing exhaust performance to increase thrust at a high Bypass Ratio (BR), and bypass throttling at low BR. In order to explore the aerodynamic characteristics of the variable injector, the unsteady motion characteristics and the congenerous thermal transport effect of three-dimensional wake vortices within the BR range of 0.18 to 0.68 are numerically analyzed. The results demonstrate that the injector can adjust the sidewall of the collapsible lobe based on the injection of bypass flow, and significantly strengthen streamwise vortices induced by lobe peak (SV-LP), which can effectively drive the migration of the heterogeneous fluid element, so as to compensate for the inefficient mixing ability of Kelvin-Helmholtz (K-H) vortices. When the BR is raised to 0.36, the increasing SV-LP will gradually replace the low-disturbance K-H vortices in the dominant position of the quasi-ordered flow field by virtue of a strong entrainment effect. Therefore, the injector can maintain a relatively high thrust augmentation in a wide working envelope. [ABSTRACT FROM AUTHOR]

This work identifies and explores several aeroacoustic metrics that allow for urban air mobility (UAM) vehicle noise prediction. An increase in production and use of UAM and distributed electric propulsion vehicles within populated civilian areas stands behind the need to minimize the noise produced by these vehicles. The Federal Aviation Administration's (FAA's) strict noise regulations on UAM aircraft compels designers to place a significant emphasis, early in the design phase, on the characterization and analysis of the external noise generated by these vehicles, namely, to ensure their design viability. To accomplish this, the present study focuses on the analysis and interpretation of predicted noise signals using a set of characteristic metrics that can be instrumental at guiding the design process. Following a thorough review of metrics standardized by the International Civil Aviation Organization as well as the FAA, seven general metrics are identified, evaluated, and discussed in the context of UAM noise prediction. When used in conjunction with a modern surface-vorticity panel code, these metrics are shown to provide an effective assortment of tools to concisely describe UAM-based acoustic signal properties. [ABSTRACT FROM AUTHOR]

Aerodynamic heating and impact resistance present significant challenges for hypersonic aircraft. This study introduces a novel multi-jet strategy to enhance both resistance reduction and thermal protection performance of hypersonic aircraft. Computational Fluid Dynamics (CFD) analysis is employed for the aerodynamic evaluation. The results demonstrate that this novel strategy effectively mitigates the shock waves, the peak pressure coefficient and Stanton number have been reduced by 64.3 % and 73.2 %, respectively. Through a comprehensive analysis of the influencing factors, it has been found that increasing the pressure ratio of the root jet significantly lowers the heat flux and pressure on the blunt body, albeit at the cost of an increased total flight resistance. When the nozzle on the side of the airway is oriented perpendicularly to the incoming flow direction, a notable reduction in resistance and aerodynamic heating on the blunt body is noted. By increasing the length-diameter ratio of the spike, a significant decrease in the pressure coefficient of the blunt body is achieved, the Stanton number remains largely unaffected. This study offers insights into the engineering application of strategies for reducing resistance and heat in hypersonic aircraft. • Propose a novel configuration that combines the passive jet with the active jet. • Conduct a numerical study to investigate the reduction of resistance and the thermal protection capability of the proposed configuration. • Analyze how the design parameters impact the performance and provide the explanation for the underlying mechanism. [ABSTRACT FROM AUTHOR]