Supersonic Vehicle Research Articles

Numerical investigations of supersonic combustion characteristics of thermally cracked fuels at typical pyrolysis temperatures

Insight into supersonic combustion characteristics of thermally cracked fuels in a scramjet combustor is crucial for the development of next-generation supersonic flight vehicles. In the present study, four typical thermally cracked fuels at pyrolysis temperatures of 450, 550, 600, and 650 ℃ (denoted as ckori, ck550, ck600, and ck650) were selected. Improved delayed detached eddy simulation (IDDES) modeling with a compact skeletal mechanism and a transported multi-regime flamelet model was carried out to simulate the supersonic combustion of the target fuels in a full-scale scramjet combustor. Supersonic flow, mixing, and combustion characteristics of these fuels were quantitatively compared. Discrepancies in the combustion process are mainly at the cavity region, where the mixing efficiencies and combustion intensities of the first three fuels increase monotonically with the increasing pyrolysis temperature, while that of ck650 locates between ck550 and ck600. The increase in pyrolysis gas content will enhance the mixing process, while the increase in injection pressure will suppress the heat/mass transport between fuel and incoming air. The mixing process dominates the distributions of intermediate radicals and main products. The first three fuels tend to combust under fuel-rich condition, while ck650 tends to combust under the stoichiometric condition, which explains the distributions of OH, CO, and CO2 concentrations. The premixed and non-premixed combustion modes coexist in the model scramjet combustor, and the non-premixed combustion mode dominates the combustion process of thermally cracked fuel.

Assessment of drag prediction techniques based on radar data for supersonic vehicles

Flight testing is probably the most accurate approach for defining the aerodynamic characteristics of a flying vehicle. This is justified by the fact that compared to other approaches, either engineering or computational ones, flight testing is indeed the one that perfectly resembles the real flight environment. Flight data are either obtained by measuring the vehicle’s kinematics via onboard sensors or by tracking the vehicle’s flight via (commonly) radars. The advantages of the latter approach are evident in cases where modifying the vehicle design is not possible or in cases where rival/enemy vehicles are examined. The key issue for this approach is the quality and demands of the technique by which vehicle aerodynamic characteristics are reduced from flight-tracked data. In the open literature, different techniques are used to analyze radar data of vehicle position and utilize them to predict vehicle drag coefficient. Each technique has its strengths and pitfalls. In this paper, the three well-used techniques of drag estimation from radar data namely, Least Square (LS), Maximum Likelihood Estimation (MLE), and Stepwise regression (SR), are considered. The underlying principle, the output, and the range of validity for each technique are addressed with emphasis on what differentiates each of them. The viability and validity of the three techniques are addressed based on the own flight testing of a free-flight supersonic vehicle and using the point-mass flight model. Meteorological data are also recorded and flight conditions are used to enhance the resulting calculations. Based on experimental data available from the literature, a comparison is conducted for the techniques examined. Considering the lack of flight data utilized, and in conjunction with data from the literature, it has been concluded that the SR technique outperforms for a higher sample rate, however, the MLE is more feasible when there is a lack of data.

Aerodynamic Effects of Surface Protuberance Size on Slender-Bodied Supersonic Vehicle

The presence of protuberances can create an asymmetric flowfield, which contributes to side forces in slender-bodied launch vehicles. In this study, we conduct numerical calculations using a supercomputer at Japan Aerospace Exploration Agency (JAXA) on a slender body with a different-sized protuberance at Mach 1.5 to systematically determine the aerodynamic effects of the protuberance size. The protuberance size is varied in its height and width. According to the results, it is demonstrated that the side force significantly increases when the height of the protuberance increases. This is because, the higher the protuberance, the farther the wake vortex produced by the protuberance moved away from the body. Consequently, the flow asymmetry between the protuberance side and clean side is augmented, and the side force increases. In contrast, the side force is almost constant when only the width of the protuberance is changed. The results of this study indicate that when attaching the protuberance to the vehicles the height of the protuberance should be lowered, and the width of the protuberance should be increased to secure the volume of the protuberance and reduce the increase in side force.

Base Flow and Drag Characteristics of a Supersonic Vehicle with Cold and Hot Jet Flows of Nozzles

Base drag has a significant effect on the overall drag of a projectile in a supersonic flow. Herein, the base drag and flow characteristics of cold and hot gas flow in a supersonic flow are analyzed via numerical simulations. The hot gas flow is simulated using a chemical equilibrium application code based on hydrogen combustion. Two types of nozzle configurations, namely conical and contoured, are chosen for the simulation. The simulation results reveal that the change in base drag is 5–85% according to the injection gases. In the over-expanded and slightly under-expanded conditions, the base drag decreases in the hot gas flow, owing to the weak expansion fan caused by the high-temperature nozzle flow expansion, whereas in the highly under-expanded condition, the base drag decreases, owing to the strong shock wave near the base caused by the deflection of the recirculation region toward the body wall. In addition, the variations in base flow structures are observed differently compared with the cold flow; for example, a weak oblique shock wave at the nozzle exit, an increase in the distance between the shock wave and base, and deflection of the recirculation region based on the body wall are observed.

Surrogate-based optimization on bump for shock wave/boundary layer interaction control

Supersonic and hypersonic vehicles have complex shock wave/boundary layer interaction problems during flight, and it adversely affect the aerodynamic performance of the vehicle. In this paper, the bump passive control approach is introduced to control the shock wave/boundary layer interaction, and its configuration is optimized with the minimum separation area and the maximum total pressure recovery coefficient. The Kriging model is established by means of the Latin hypercube experimental design approach to obtain the sample points. The Non-dominated Sorting Genetic Algorithm-II (NSGA-II) algorithm is utilized to optimize the parameters of the bump configuration, and the flow field structures of the uncontrolled condition, the special basic condition and the optimized condition are compared and analyzed. The obtained results show that the surrogate-based optimized model shows better performances than the basic condition. In the parameter setting range considered in this study, the optimized Case 3c with the middle position in the optimization front owns a larger total pressure recovery coefficient and a smaller separation area, with its separation area and total pressure recovery coefficient being 0.80 mm2 and 67.39% respectively.

Numerical study of heat transfer and load-bearing performances of corrugated sandwich structure with open-cell metal foam

Critical engineering applications require lightweight components that can effectively dissipate heat and provide sufficient load-bearing capacity. Examples of such applications include the jet blast deflector (JBD) on an aircraft carrier and engine combustion chambers in supersonic vehicles. While lightweight corrugated sandwich structures (CSSs) exhibit excellent load-bearing capacity, their heat transfer capacity falls short. To overcome this limitation, a strategy of filling the CSS with an open-cell metal foam (MF) is proposed to create a sandwich structure with ultralight load bearing and heat transfer capabilities. Numerical simulations were conducted to study the heat transfer and load-bearing performance of this new structure. The findings indicate that the use of a MF significantly improves the heat transfer capabilities of the CSS. Compared to the CSS, the Nusselt number of MF and CSS-foam composite at a Reynolds number of 240 were enhanced by 51.3% and 102.3%, respectively. The overall thermal performance of CSS-foam composites was optimized under the same pumping power constraints. In turbulent conditions, the overall thermal performance of CSS-foam composites was 5.9% to 55.4% higher than that of CSS in turbulent conditions. Furthermore, the simulations showed that when the CSS was subjected to static and quasi-static compression, the maximum von Mises stress occurred at the connection between the panel and the web.

Angular rate enhanced Overload Controller Design for Low-speed UAVs

This paper investigates the overload controller for low-speed UAV maneuvering flight. Compared with high-speed and supersonic vehicles, low-speed UAVs are more sensitive to atmospheric disturbances during maneuvering and their signals are more sensitive. This paper combines UAV overload control theory and L1 adaptive theory to propose a longitudinal attitude control method based on overload control for low-speed UAVs. The method uses overload as the outer loop and pitch angle rate as the inner loop, and provides a reliable control scheme for maneuvering low-speed UAVs. Through simulation experiments, the controller proposed in this paper is compared with the conventional classical three-loop overload controller. The results show that the controller of this paper has faster angular rate response and stronger anti-interference capability under the same overload corresponding conditions.

Preliminary design of next generation Mach 1.6 supersonic business jets to investigate landing & take-off (LTO) noise and emissions – SENECA.

With the approach of next generation supersonic transport entry into service, new research activities were initiated to support updates on ICAO regulations and certification processes for supersonic transport vehicles. Within this context, the EU Horizon 2020 SENECA project has been launched to investigate the levels of noise and gaseous emissions in the vicinity of airports as well as the global climate impact of next generation supersonic civil aircraft. This paper introduces some of the preliminary outcomes of this investigation. It presents the preliminary design and performance analysis of a Mach 1.6 business jet, following an integrated aircraft-engine design approach. The preliminary design was performed accounting for the limitations posed by future environmental restrictions on respective subsonic vehicles. The market space and mission route definition exercise assumed only “over-sea” supersonic operations, while for “over-land”, only subsonic operations where allowed. Parametric studies on engine integrated design demonstrated modest core temperatures while cruising and the significant impact of engine installation on performance. At this first design iteration, assuming current state of the art technology, the Mach 1.6 business jet showed good potential to satisfy the predicted mission requirements while respecting the environmental constraints in terms of Landing & Take-Off (LTO) noise and emissions.

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

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

Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

Supersonic shock wave boundary layer interactions are common to inlet flows of supersonic and hypersonic vehicles. This paper reports on wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 turbulent shock wave boundary layer interaction experiment at the NASA Glenn Research Center. The boundary layer upstream of the interaction was nominally axisymmetric and two-dimensional. A conical centerbody with a 16 deg half-angle and a maximum radius of 0.147D of the test section diameter was employed to generate a conical shock wave, where D is the test section diameter. Asymmetric (swept) interactions were obtained by displacing the shock generator away from the test section centerline. The present simulation is for a shock generator displacement of D/6. Results from the asymmetric simulation are compared with results from an earlier simulation of a corresponding axisymmetric interaction. The experimental Reynolds number based on test section diameter was ReD=4×106. For the simulations, the Reynolds number was lowered to ReD=4×105 to keep the computational expense of the simulations within limits. Compared to the axisymmetric interaction, the streamwise extent of the separation varies considerably in the azimuthal direction for the asymmetric interaction. The separation is strongest at the azimuthal location that is closest to the shock generator. The streamwise extent of the separated flow regions is noticeably reduced and substantial crossflow is observed between the locations that are closest and farthest from the shock generator. A Fourier analysis of the unsteady flow data indicates low-frequency content for the separated region that is closest to the shock generator. Away from this region, with increasing sweep angle and cross-flow, the low-frequency content is diminished. A proper orthogonal decomposition captures spanwise coherent structures for the more two-dimensional parts of the interaction.

Study on aerodynamics characteristics of the effect of leading-edge radius of FDM printed (ABSi material) supersonic aerofoil

Currently, supersonic vehicles are studied as future civil transports. The selection and design of aerofoils plays an important step in the aircraft design process. NACA 64-008A aerofoil is chosen for the investigation which is produced using ABSi (Acrylonitrile butadiene styrene) by Fused Deposition Modelling (FDM). The focus of this work is to investigate numerically the aerodynamic characteristics of supersonic aerofoil, NACA 64-008A, for the effect of varying the leading-edge radii along with variations in angle of attacks. The simulation is carried out using commercial CFD code ANSYS Fluent. The closure for the Reynolds Averaged Navier Stokes equations is achieved by using the k-ω SST turbulence model. The coefficients of lift and drag for three distinct leading-edge radii, namely 4.8 mm, 5 mm, and 5.2 mm at Mach 2.25 with a varied angle of attack (-5°to 5°) are investigated. k-ω SST turbulence model is selected to deal with the flow turbulence and achieve the closure for Reynolds averaged Navier Stokes equations. The CL/CD ratio increases as the leading radius increases. The maximum value of CL/CD ratio is observed for 5.2 mm leading edge radius at an angle of attack, α = 5°. Moreover, the drag decreases with increase in lift values. It is clearly seen that increase in the leading-edge radii of the supersonic profile has a considerable effect in the increase in coefficient of lift. Thus, the variations in the leading-edge radii of the supersonic aerofoil enhances the aerodynamical characteristics.

High-Performance Attitude Control Design of Supersonic Tailless Aircraft: A Cascaded Disturbance Rejection Approach

This paper focuses on the triaxial augmentation ability of the active disturbance rejection control (ADRC) technique on the tailless layout with a fully moving wing tip to achieve high control performance for the supersonic tailless aircraft. Firstly, the stability characteristics and controllability of the flying wing layout are analyzed to determine the coupling characteristics of this kind of aircraft. Secondly, an attitude controller is designed based on ADRC theory, and the linear ADRC frequency domain analysis method is introduced to analyze the influence of the bandwidth of linear extended stator on the control system. In addition, the tuning process of the attitude control law is given. Carrier dropping simulations of flight missions under nominal condition, model parameter perturbation, and wind disturbance are conducted. The results show that the designed controller can achieve full-speed domain triaxial augmentation of supersonic flying wing. This work has the potential to significantly boost the engineering acceptability and robustness of supersonic aircraft control design in real-world scenarios. The presented cascaded ADRC approach can significantly improve the performance and robustness of supersonic vehicles.

Mechanical response and damage mechanism of C/SiC composites impacted by high-velocity water jet

Rain erosion is a potential hazard for supersonic vehicles, with severe damage to materials that may be impacted by raindrops. In this paper, a series of impact tests of 413–572 m/s are carried out on a 3 mm-thick 2D C/SiC composite specimen using a single impact waterjet apparatus. The typical morphology of C/SiC specimen is obtained by single jet impact test. Under the multi-drop impact, the stress wave interaction is enhanced, and the internal damage of the specimen is severe, showing a funnel-shaped damage. Moreover, the C/SiC specimen is penetrated after 5 drops of impact. Quasi-static tensile tests were employed to quantify the post-impact strength of the specimen, during which the digital image correlation (DIC) method was used to obtain the strain value, at the same time acoustic emission (AE) signal was detected and processed by the K-Means to reveal the damage evolution.

Sonic Boom over land avoidance and the impact on economic feasibility

There has been renewed interest in commercial supersonic air travel in recent years with a number of new efforts being pursued by several companies. They currently advertise entry into service of these new concepts within the next decade. Current regulations do not permit overland operations for supersonic vehicles due to the sonic booms produced during flight. In order to capture a favorable market share, these vehicles will need to serve routes while abiding with the over land boom restrictions. This study presents a comprehensive approach to overland-prohibited supersonic flight routing and route-specific demand forecasting for commercial supersonic air travel, in order to assess the expected market. This also includes the need for the seasonally dependent need to avoid secondary sonic boom exposure of land.

Nonlinear Aeroelastic Characteristics of a Supersonic All-Movable Fin with Single or Multiple Freeplay Nonlinearities

Establishing the aeroelastic characteristic of all-movable fins with freeplay nonlinearities is one of the most common problems in the design of supersonic flight vehicles. In this context, this study provided novel points of view on the nonlinear aeroelastic characteristics of an all-movable fin with freeplay nonlinearities in its root. The unsteady aerodynamic model that was employed uses the second-order piston theory considering thickness effects. For a system with multiple freeplay nonlinearities, a discrete scanning method based on the describing function method was established to solve the limit-cycle oscillations (LCOs) and avoid the loss of solutions. Combining this with the time-domain integration method, the influences of the support stiffness at the root of the fin and the freeplay size ratio of the bending and torsional degrees of freedom on the dynamical response of the system were analyzed. The results demonstrate that systems with a single freeplay nonlinearity exhibit two completely different types of LCO, while systems with multiple freeplay nonlinearities exhibit complex dynamical behaviors such as LCO and quasi-periodic and chaotic motions. The path of a quasi-periodic torus breaking into chaos was observed. Furthermore, a harmonic initial condition for the time-domain integration is proposed; this can be used for a quick check of the frequency-domain calculation results.

Thermal insulation and structural reliability of modified epoxy resin‐based ablation thermal protection coatings in aerothermal‐vibration coupling environment

AbstractNovel hybrid ablation thermal protection coatings (FHMP‐ATPCs), employing iron trioxide (Fe2O3) powder, hollow glass microspheres, and mica powder as the fillers in hydroxyl‐terminated silicone oligomer‐bridged epoxy resins (PSG) copolymer, is investigated using an aerothermal‐vibration coupling test system. The ablation behavior and structural reliability of FHMP‐ATPCs with varying coating thickness were studied. During the test, the total enthalpy of airflow and dynamic pressures are 23 MJ/kg and 300 Pa, accompanied by the random vibration with a frequency of 20–2000 Hz and a total root‐mean‐square acceleration of 14.9g. The maximum surface and back‐face temperatures of the coating with the thickness of 2 mm reached 836.2°C and 156.4°C, respectively. Results also showed that the reduction of thickness obviously suppressed the surface temperature and increase in back‐face temperature yet maintaining high structural reliability. Compared with DGEBA‐based coatings, the PSG‐based coatings showed excellent structural reliability during the test. The study provides a solution for obtaining high performance ATPCs, which are highly desired for supersonic vehicles.

Aeroelastic tailoring of composite rudder skin considering variable angle tow laminates by a hybrid backtracking search-JAYA-Sine Cosine Algorithm

Flutter is a typical dynamic instability phenomenon of the wing or rudder structure. Besides, the phenomenon of aerodynamic heating will be quite significant and temperature can rise rapidly with the increase of the flight speed, thermal flutter characteristics should be taken into consideration for the rudder structure of supersonic vehicle. To improve the thermal flutter characteristics of the rudder structure and meet the design requirements of a lightweight structure, this article optimizes the variable stiffness laminates of the skin on the premise of keeping the thickness of the skin unchanged, so as to improve the critical flutter speed of the rudder. A novel intelligent optimization algorithm named BJASCA, which is a hybrid algorithm coupling of Backtracking Search Algorithm, Sine Cosine Algorithm, and JAYA Algorithms is proposed in this article to improve the efficiency of structural optimization and this algorithm shows obvious advantages in searching global optima compared with several other algorithms. The finite element model and aerodynamic model of the rudder are established to calculate the critical flutter speed. The results represent that the rudder structure optimized by the BJASCA algorithm shows better flutter characteristics and the critical flutter speed can be increased by 76.4% compared with the un-optimized structure.

Research on the influence of aero-optical effects on starlight atmospheric refraction navigation under supersonic conditions

In the starlight atmospheric refraction navigation when the starlight transmits in the supersonic flow field, the aero-optical effect will reduce the accuracy of navigation. In this paper, the aircraft model is established by ICEM and Fluent is used to simulate the atmosphere density distribution at different altitudes and speeds. Then, the principle of geometric optics is used to track the starlight, the angular deviation of starlight transmission is deduced, and finally, the influence of different speeds and altitudes on starlight atmospheric refraction navigation is analyzed. The results show that the aero-optical effect produced by supersonic vehicles is related to the flight altitude and flight speed. Taking the flight altitude of 20 and 30 km as an example, when the flight speed is Mach 2, the angular deviation caused by the aero-optical effect is 1.045 and 0.699“ respectively, and when the flight speed is Mach 10, the angular deviation is 20.075 and 4.643”, respectively. Therefore, the aero-optical effect can be ignored at the altitude of 30 km and above. However, the influence of the aero-optical effect at 20 km needs to be judged according to the flight speed.

Multi-scale entropy based damage detection for thermal protection structures with variation of ambient temperature

The thermal protection structures of supersonic vehicles are vulnerable to damage in the extreme environment of high temperature. With advantages of the large propagation range and high sensitivity to damage, guided waves show great potential for structural health monitoring. However, guided waves are susceptible to ambient temperature change, resulting in low reliability of damage detection results. Therefore, this paper proposed a multi-scale entropy feature extraction method for structural damage detection. Firstly, to eliminate the influence of multi-mode guided waves, a sliding window was used to extract the wave packet of different signals to analyze disturbance caused by damage and temperature. Then, an ant colony optimization algorithm was introduced as a feature fusion method to improve the classification performance. To evaluate the performance of features selected by the ant colony optimization algorithm, K-means clustering algorithm, and silhouette coefficients were utilized to calculate the evaluation function and represent the characteristics of damage. Finally, a set of guided wave signals from 20°C to 40°C were obtained to investigate the influence of temperature variation on the damage feature extraction. The results show that the multi-scale entropy method based on the sliding window can effectively extract damage features in the condition of temperature change.

Thermal stress stability of hydrocarbon fuels under supercritical environments

This study investigates the thermal stressing characteristics of hydrocarbon fuels to develop thermally stable fuels for supersonic vehicle applications. The thermal stability characteristics of two multicomponent hydrocarbon fuels, namely F-1 and F-2 fuels, under subcritical and supercritical conditions, are examined for a wide range of temperatures (30–500 °C) and pressures (15–50 bar) using a semi-batch reactor. Various analysis techniques, such as UV-Vis spectrophotometry, ASTM D86 distillation, gas chromatography, scanning electron microscopy (SEM), etc., are used to characterize the feed and product properties. The study showed that the fuels are reasonably stable up to around 400 °C and above 40 bar pressure. The stability of the fuels is relatively less under subcritical pressure at elevated temperatures. The amount of gum formation increased with the increase of both temperature and pressure parameters. Further, the SEM analysis showed oval-shaped and ribbon-like structures in the solid deposits. The present investigation may be useful in developing an appropriate fuel for regenerative cooling in supersonic applications.