Forebody Configuration Research Articles

The Role of Forebody Topology on Aerodynamics and Aeroacoustics Characteristics of Squareback Vehicles using Computational Aeroacoustics (CAA)

This study investigates the influence of forebody configuration on aerodynamic noise generation and radiation in standard squareback vehicles, employing a hybrid computational aeroacoustics approach. Initially, a widely used standard squareback body is employed to establish grid-independent solutions and validate the applied methodology against previously published experimental data. Six distinct configurations are examined, consisting of three bodies with A-pillars and three without A-pillars. Throughout these configurations, the reference area, length, and height remain consistent, while systematic alterations to the forebody are implemented. The findings reveal that changes in the forebody design exert a substantial influence on both the overall aerodynamics and aeroacoustics performance of the vehicle. Notably, bodies without A-pillars exhibit a significant reduction in downforce compared to their A-pillar counterparts. For all configurations, the flow characteristics around the side-view mirror and the side window exhibit an asymmetrical horseshoe vortex with high-intensity pressure fluctuations, primarily within the confines of this vortex and the mirror wake. Side windows on bodies with A-pillars experience more pronounced pressure fluctuations, rendering these configurations distinctly impactful in terms of radiated noise. However, despite forebody-induced variations in pressure fluctuations impacting the side window and side-view mirror, the fundamental structure of the radiated noise remains relatively consistent. The noise pattern transitions from a cardioid-like shape to a monopole-like pattern as the probing distance from the vehicle increases.

Investigating the effect of the flow direction on heat transfer and energy harvesting from induced vibration in a heated semi-circular cylinder

The present study examines the flow and heat around low-mass arched forebodies (AF), flat forebodies (FF) and a circular cylinder (CC) in the presence of buoyancy. The Richardson number prescribed in this investigation varies from 0 to 1.5 for the Prandtl number of 0.71 and Reynolds numbers of 100 and 200. Selecting five semi-circular cylinders with various length-to-diameter ratios (L* = L/D) aims to investigate the influence of L* on energy harvesting. The arched and flat surfaces exposed to the incoming flow can result in experiencing heat transfer augmentation due to the separating shear layers. The functional dependence of drag coefficient, lift coefficient, non-dimensional maximum vibration amplitude, and output power are analyzed and examined in detail. The main novelty introduced in current investigation is to determine a bluff body that generates a substantial amount of vibration in order to capture maximum energy. Another connotation of this study regards the possibility of employing a numerical approach to combine the effects of geometry and buoyancy to alter heat transfer as well as extracted flow energy. Increasing the Richardson number results in increased heat transfer and energy harvesting. This is because the interaction between fluid and bluff body in the presence of buoyancy can cause disruption of the thermal boundary layer, thereby augmenting heat transfer. In the absence of buoyancy force, the circular cylinders with L* of 0.5 and 0.6 for the flat forebody configuration gain the highest extracted power at Reynolds numbers of 100 and 200, respectively, resulting in 3.5 and 5.3 times more extracted power compared to the circular cylinder. Considering extracted power, it is shown that a D-shaped configuration is less dependent on buoyancy force compared to an inverted D-Shaped configuration. The flat forebody configuration always harvests more energy than the same body with opposite flow direction. These results are capable of providing deeper insight into configuring the thermal systems, which are able to dissipate heat loads as well as produce maximum power. Having compared some of the outcomes of present work with other studies, a good agreement has been achieved to validate the simulations.

Shock-related unsteadiness of axisymmetric spiked bodies in supersonic flow

Shock-related unsteadiness over axisymmetric spiked body configurations is experimentally investigated at a freestream supersonic Mach number of 2.0 at 0 $$^\circ $$ angle of attack. Three different forebody configurations mounted with a sharp spike-tip ranging from blunt to streamlined (flat-face, hemispherical, and elliptical) are considered. Steady and unsteady pressure measurements, short-exposure high-speed shadowgraphy, shock footprint analysis from $$x-t$$ plots, and identification of dominant spatiotemporal modes through modal analysis are carried out to explain the unsteady flow physics. The present investigation tools are validated against the well-known events of ‘pulsation’ corresponding to the flat-face case. The hemispherical case is characterized by the formation of a separated free shear layer and associated localized shock oscillations. The cycle of charging and ejection of fluid mass from the recirculation zone is identified to drive the flow unsteadiness. Such an event triggers the movement of the separated and reattachment shocks in the opposite direction with respect to each other (out-of-phase shocks motion). In the elliptical case, the overall flow field resembles that of the hemispherical case, except with dampened unsteadiness. The value of the cone angle ( $$\lambda $$ ) associated with the recirculation region is found responsible for the fluctuations caused by the charging and ejection of fluid mass. Thereby, it controls the extent of out-of-phase shocks motion. In the elliptical case, shock unsteadiness is reduced as $$\lambda $$ is observed to be smaller. Based on the gathered results and understanding, the reduction in unsteadiness associated with the aerodisk mounted on the hemispherical forebody is demonstrated and explained via the almost complete elimination of the out-of-phase shocks motion.

On the effects of orientation on flow and heat transfer from a semi-circular cylinder near a stationary wall

Numerical analysis of flow and forced convection heat transport from a two-dimensional semi-circular cylinder placed near a stationary wall is presented. The uniformly heated cylinder is within the boundary layer of the plane wall and is subjected to linear shear flow. An unsteady flow of incompressible Newtonian fluid around the cylinder is studied assuming constant properties. Two different orientations of the cylinder, i.e. the curved and flat surfaces exposed to the incoming flow, are considered, and the results are compared with those of a circular cylinder. The cylinder-to-wall gap ratio (G/D.) values are varied between 0.1 and 0.5, and the Reynolds number range of 100 ≤ Re ≤ 200 is considered. Numerical experiments are performed using Ansys Fluent. Results suggest that the presence of a stationary wall influences both the flow and thermal fields significantly. A decrement in the drag coefficient and average Nusselt number values is reported, while the lift coefficient and Strouhal number increase, resulting from shear layers’ interaction in the gap between cylinder and wall. The heat transfer enhancement ratio is higher for the flow past curved forebody than the flat forebody configuration and circular cylinder. Mean Nusselt number correlations for the range of parameters considered in the present study are provided.

Theoretical Review on Hypersonic Cruise and Acceleration Vehicle's Forebody Configuration

Theoretical Review on Hypersonic Cruise and Acceleration Vehicle's Forebody Configuration

Transonic Aerodynamic Analysis of Different Forebody Shapes of Aerospace Vehicles

One of the most important features of an aerospace vehicle design is its shape, especially the forebody part, since it defines the main characteristics of the vehicle aerodynamic. In this context, the present work has the purpose of investigating the transonic aerodynamic regime in typical vehicle models, with different frontal geometry configurations, using the experimental techniques pressure sensitive paint (PSP) and Schlieren. The aim is to get more insights about the flow patterns over the selected models, regarding the transonic regime, in terms of qualitative density field evaluation, and the identification of important aerodynamic parameters, such as shock wave positioning and pressure gradients over the model, as well as the influence of the forebody configuration in the flow pattern. The test campaign was carried out in the Pilot Transonic Wind Tunnel of the Institute of Aeronautics and Space (IAE; in Portuguese, Instituto de Aeronáutica e Espaço), varying Mach numbers from subsonic to low supersonic. Five forebody configurations mounted in a cylindrical body having the standard cone-cylinder shape as baseline was considered. A description of the phenomena occurring around and on the models surface for all five models is presented, through analyses of the Schlieren images and PSP results.

Parameterization and optimization for the axisymmetric forebody of hypersonic vehicle

The axisymmetric hypersonic vehicle has attracted an increasing attention due to its simple structure and high volumetric efficiency. In order to find the optimal aerodynamic configuration of the axisymmetric hypersonic forebody, a novel parameterization method based on the theory of Non-uniform rational B-spline (NUBRS) curve and conical flows is proposed. Then, the optimization work which couples the global optimization algorithms NSGA-II with the computational fluid dynamics (CFD) fluid analysis is carried out. The optimization results show that configurations with higher lift-to-drag ratio as well as higher volumetric efficiency than the commonly used Von Karman Ogive configuration can be found. By analyzing flow structures of three different representative configurations, the reasons that lead to the different lift-to-drag ratios are found. Finally, the effects of angle of attack on the lift-to-drag ratio of three representative forebody configurations is studied.

Design of hypersonic forebody with submerged bump

A double shock waverider forebody configuration, with curved surfaces and known pressure fields and shock arrays, is constructed by a stream-tracing approach. The compression surface consists of a wedge and conical shocks. The conical shock results from a modified wave-derived bump surface that diverts the boundary layer before the inlet entrance. The design is fully computational fluid dynamics based and emphasis is placed on the compact design with boundary layer diverting ability. Controlling or diverting the thick boundary layer safely is a difficult challenge in hypersonic flight vehicle design especially when the inlets are highly integrated with the fuselage. Numerical simulations show that the new combination can divert a significant fraction of boundary layer before the inlet and maintains a good compression ratio for propulsion efficiency at Mach 5.0. Effects of forebody aerodynamics on the integrated inlet and comparisons with other systems are described in this paper.

A modular RANS approach for modeling hypersonic flow transition on a scramjet-forebody configuration

In this study, the laminar-turbulent flow transition scenario on a 20%-scale X-51A forebody configuration, is simulated with a recently proposed RANS (Reynolds-Averaged Navier–Stokes) transition/turbulence model, which takes into account of the rational effects of compressibility, crossflow and flow-separation characteristics on different instability modes in 3-D boundary-layer flows (Fu and Wang 2013). First, we perform model validation with a number of available experiments on boundary layer transition including hypersonic flows past sharp leading-edge flat plates and cones, nose blunted cones and double ramps. Moreover, a new formula is established to convert measured free-stream noise level to FSTI (free-stream turbulence intensity) that is more adopted by engineering transition approaches. The results show generally good agreement with experimental data, proving that the present model responds accurately to both FSTI and leading edge radii. Then, a simplified analysis is proposed to characterize the present model behavior in pre-transitional flow regions. The results identify the model capability of the transition onset prediction with regard to the influences of nose bluntness as well as roughness elements. Finally, the success in its application to the forebody configuration (with/without trips) under both quiet and noisy conditions, indicates that the mixed-mode transition scenario indeed benefits from such a modular prediction approach, which embodies current conceptual understanding of the transition process.

Flow mechanism of self-induced reversed limit-cycle wing rock for a chined forebody configuration

The wing rock phenomenon reduces the maneuverability and affects the flight safety of modern advanced fighters, such as the F-35, which have chined forebodies. Understanding the flow mechanism is critical to suppressing this phenomenon. In this study, experiments were conducted to reveal the motion and flow behavior over a chined forebody configuration. The tests were performed in a wind tunnel at an angle of attack of [Formula: see text] with a Reynolds number of [Formula: see text]. Reversed limit-cycle oscillation was discovered in the free-to-roll tests. The unstable rolling moment around zero roll angle in the static case suggests that the model tends to be driven away from zero roll angle. Thus, the model cannot maintain its equilibrium at zero roll angle during free-to-roll motion. The unstable rolling moment is generated by the wing vortex structure above the upward wing, which is induced by the forebody asymmetric vortices. During wing rock, the wing vortex structure appears above the upward wing at a large roll angle after crossing zero roll angle owing to a time lag in the forebody vortex position, which is conducive to the motion. The forebody asymmetric vortices are thus the key to induce and maintain the motion.

An experimental study on chaotic oscillation of a chined forebody configuration in roll

The chaotic oscillation is an aperiodic oscillation rolling between a positive and a negative equilibrium positions with a chaotic behavior. This oscillation pattern was investigated over a chined forebody configuration through a series of wind tunnel tests at an angle of attack of 40° with a Reynolds number of 1.87 × 105. The wing vortex flow around zero roll angle induced by the forebody asymmetric vortices flow is crucial to the motion. Under static conditions, the strong leeward wing vortex causes the unstable lateral stability around zero roll angle and leads to the formation of a positive and a negative stable equilibrium position of the rolling moment, which is considered as the premise of forming the motion. The strong leeward wing vortex at a low angular velocity can stop the model from crossing zero roll angle and drive the model to increase amplitude, which is the key to explain the flow mechanism. Weakening of the leeward wing vortex due to a prominent hysteresis at a high angular velocity make the model cross zero roll angle, but the appearance of the strong windward vortex after crossing zero roll angle decelerates the rolling motion and makes the amplitude decrease.

Experimental research on a hypersonic configuration with blunt forebody edges

A design method, for hypersonic forebody configuration with blunt edges, was developed by lift body and waverider. Experimental study was conducted. It was showed that the difference of pressure coefficient, at centre-line and side-line on forebody precompression surface, was between 2.5% and 16.7% within the Mach number 5 to 7 and attack angle 0 to 12. These differences were slightly bigger than those of non-blunt-edges forebody. However, experimental data also show that the present configuration can achieve the closed shockwave, and can also realize the automatic overflow of boundary layer on forebody pre-compression surface. It is also observed that the present forebody gets a good aerodynamic performance at high Mach numbers. But with increasing attack angle, the advantage was weakened gradually. And, at the degree of side-slip exceeding 8, the side-line pressure is higher than that of center-line, which indicates that the forebody boundary layer does not achiev automatically overflowing to side surface.

Numerical Study on Aerodynamic Design of Hypersonic Vehicle Forebody

The lift forebody configuration of a small hypersonic vehicle is designed by using the wedge angle method in this paper. The lift forebody created has been optimized by the simplex method with a penalty function. The aerodynamic characteristics of the forebody optimized are investigated by numerical method. The research results show that the wedge angle method is a high efficient way to generate the lift forebody of the hypersonic vehicle; On the design mach number, there is pressure leaking between the upper and lower surface of lift forebody, which leads to lateral flow in the spanwise on the precompression plane, and which create the unhomogeneity of inlet flow field; Adding side skirts on the both sides, which can reduce the lateral flow on the forebody's precompression plane, it can raise the forebody lift.

A TIPSO algorithm assessment for aerothermodynamic optimization of hypersonic compression systems

A two-step improved particle swarm optimization (TIPSO) algorithm was recently proposed and used for the optimization of flexible satellite-control parameters. It was found to be much more stable and much less complex than other evolutionary algorithms. In this article the efficacy of the TIPSO algorithm is investigated for multidisciplinary optimization of aerothermodynamic parameters of performance, cowl deflection angle and shock–boundary layer separation on a cone-derived, wedge-integrated hypersonic (waverider) compression system. This algorithm uses an aggregate objective function. Optimization results from the TIPSO algorithm are compared with those obtained from a hybrid genetic algorithm, particle swarm optimization using an inertial weight approach and a multi-objective genetic algorithm. Since optimality of the forebody configuration is the basic requirement, the optimization variables selected are the isolator exit Mach number, static pressure ratio across the forebody–inlet configuration, cycle temperature ratio and cowl (inlet) deflection angle, which is the maximum deflection angle corresponding to the optimum oblique shock angle generated by the inlet cowl. Simulation results show that for the given case study only the TIPSO algorithm can locate the globally best aerothermodynamic–geometric parameters, compared with all the other optimization methods.

超音速機のエンジン／機体統合化に伴う空力的影響について

In the present study, the influence of the engine unit (air-intake, nacelle and nozzle) to the aerodynamic characteristics of the airframe was experimentally investigated at a free stream Mach number of 2.75. A series of wind tunnel tests was performed in NAL's M4 supersonic wind tunnel for the airbreathing engine unit/airframe integrated models of which the engine unit was installed on the bottom surface of the fuselage. These models were designed for Mach 3 flight. To identify the flow field (e. g., boundary layer and vortices) upstream of the engine unit, pitot pressure survey and flow visualizations by schlieren and oil-flow techniques were applied. The streamwise vortices due to down wash from upper side to lower side of forebody were observed at the low angle of attack, but these vortices could be suppressed by the modification of the forebody configuration. The diverter worked effectively so that the upstream boundary layer was removed. With the measurements of aerodynamic forces and schlieren visualization, it was found that the flow state, i. e., start, buzz or unstart condition, of air-intake affects the upstream flow field and the aerodynamic characteristics of the airframe. In the case that the shock spilled out from the air-intake, the aerodynamic force normal to the airframe axis decreased as the angle of attack increasing, because the force in the nozzle decreased due to higher loss of the total pressure in the engine unit.

Aerodynamic characteristics of delta wing–body combinations at high angles of attack

AbstractAn experimental study investigating the aerodynamic characteristics of generic delta wing-body combinations up to high angles of attack was carried out at a subsonic Mach number. Three delta wings having sharp leading edges and sweep angles of 50°, 60° and 70° were tested with two forebody configurations providing a variation of the nose fineness ratio. Measurements made included six-component forces and moments, limited static pressures on the wing lee-side and surface flow visualisation studies. The results showed symmetric flow features up to an incidence of about 25°, beyond which significant asymmetry was evident due to wing vortex breakdown, forebody vortex asymmetry or both. At higher incidence, varying degrees of forebody-wing vortex interaction effects were seen in the mean loads, which depended on the wing sweep and the nose fineness ratio. The vortex breakdown on these wings was found to be a gradual process, as implied by the wing pressures and the mean aerodynamic loads. Effects of forebody vortex asymmetry on the wing-body aerodynamics have also been assessed. Comparison of Datcom estimates with experimental data of longitudinal aerodynamic characteristics on all three wing-body combinations indicated good agreement in the symmetric flow regime.

Application of a full potential method to supersonic aircraft designand analysis

A method based on the conservation form of the full potential equation has been used to analyze realistic aircraft configurations at supersonic speeds. A fighter forebody with and without a canopy and a supersonic cruise wingbody concept have been addressed in the Mach 1.41-2.30 speed range. Comparisons of predicted and measured surface pressure distributions and lift and drag for the forebody configurations showed excellent to good correlations, although some oscillations hi the computed pressures were observed. Good to excellent results for the wingbody configuration were obtained as well. The nonlinear behavior of the pitching moment was well predicted, although the magnitudes were somewhat low. These analyses have all been conducted using single-precision arithmetic on a 32-bit minicomputer. Execution times averaged 30-45 min for the forebody configurations and slightly over 4 h for the wingbody.

Bus drag reduction by the trapped vortex concept for a single bus and two buses in tandem

A wind-tunnel investigation of a 1/20 scale model of a modified Denning Mono Mark II bus indicates that the drag coefficient can be reduced to 0.29 by using optimal radii on the forebody/roof junction, 8° tapers behind the forebody and in front of the tail and incorporating a vortex trapping cavity in the tail. Surface pressure distributions and surface-mounted tufts on two forebodies and two tails provide the detailed flow behaviour. For the vortex cavity tail a vortex is well-defined over the top half of the tail but the underbody upwash interferes with the vortex flow over the lower half. The flow field between two in-line buses is determined at various gaps up to 0.472 D ( D = rmhydraulic diameter of the base) for two tails and two forebody configurations. An forebdoy/tail combinations produce a thrust ttail / D, T =2.087 tail is mode iffective in increasing the tail thrust than a flat tail, particularly at larger gaps (⩾ 0.314 D). Either tail, used with a blunt forebody, is effective in reducing the single-bus forebody drag coefficient of C D,F = 0.172 to 0.084–0.095 for gaps ⩽ 0.157 D.

Side-Force Alleviation on Slender, Pointed Forebodies at High Angles of Attack

A new device was proposed for alleviating high angle-of-attack side force on slender, pointed forebodies. A symmetrical pair of separation strips in the form of helical ridges are applied to the forebody to disrupt the primary lee-side vortices and thereby avoid the instability that produces vortex asymmetry. Preliminary wind tunnel tests at Mach 0.3 and Reynolds no. 5,250,000 on a variety of forebody configurations and on a wing-body combination at angles of attack up to 56 degrees, demonstrated the effectiveness of the device.