Search

An evolutionary algorithm (EA) was applied in this study to optimize the landing flight path of a delta-winged supersonic transport (SST). However, it is difficult for a delta wing with a large sweepback angle to reduce the aerodynamic drag during supersonic cruising to gain sufficient lift force at low speeds, particularly during takeoff and landing. Besides, high-fidelity computational fluid dynamics is required to evaluate the flight path with a complex flowfield. This study performed an efficient flight simulation based on the Kriging model-assisted aerodynamic estimation to carry out global optimization. Then, the designs of the flight and control sequence were realized for time-series optimization of effective SST landing. To develop the EA, two design scenarios were considered; one involved only the elevator, which is an aerodynamic control surface that controls the aircraft, and the other involved introducing thrust control in addition to elevator control. In the scenario involving only elevator control, feasible solutions could not be obtained owing to the poor low-speed aerodynamic performance of the SST. This paper presents several feasible solutions enabling reasonable SST landing performance in the scenario involving the elevator and thrust controls along with descriptions regarding the optimum flight and control sequences. In addition, we analyzed the solutions by analyzing the variance to obtain qualitative information. Consequently, we determined that elevator control was considerably effective in cases with the microburst effect than in cases without the microburst effect.

An engineering model of a plug nozzle is developed. The nozzle is intended for use in the engine of a prospective supersonic civilian airplane. On the basis of the model, two plug profiles are obtained for different nozzle configurations. The thrust is also estimated, taking account of the internal losses and the external aerodynamic drag of the casing. Preliminary estimates of the loss of thrust are made, taking account of the same factors. The proposed model and the results obtained may be useful in selecting the best engine structure and in optimizing the control program for the power unit.

The aerodynamic performances and flow features of the capsule/rigid disk-gap-band (DGB) parachute system from the Mach number 1.8 to 2.2 are studied. We use the adaptive mesh refinement(AMR), the hybrid tuned center-difference and weighted essentially non-oscillatory (TCD-WENO) scheme, and the large-eddy simulation (LES) with the stretched-vortex subgrid model. The simulations reproduce complex interaction of the flow structures, including turbulent wakes and bow shocks, as well as bow shocks and expansion waves. The results show that the calculated aerodynamic drag coefficient agrees well with the previou simulation. Both the aerodynamic drag coefficient and the aerodynamic drag oscillation of the parachute system decrease with the increase of the initial Mach number of the fluid. It is found that the position and angle of the bow shock ahead of the canopy change as the Mach number increases, which makes the flow inside the canopy and the turbulent wake behind the canopy more complex and unstable.

The aerodynamics of re-entry vehicles has always been in the forefront among space research. With the increasing interests in space exploration, it is important to understand the aerodynamics of spacecrafts and identify an optimum design to overcome the forces due to the earth's atmosphere. In this paper, Computational Fluid Dynamic (CFD) analysis has been carried out on ballistic bodies with a single, double and triple aero-disk spikes. The study also involves the presence and absence of a blunt tip to aid the analysis of flow behaviour on these bodies under two scenarios namely zero-degree angle of attack and at supersonic speed (Mach 2.0). The aerodynamic drag and surface heat transfer effects of these spikes on three different blunt body configurations has been studied. The shape of the blunt bodies vary based on the nose bluntness factor. The flow patterns of these configurations were three-dimensionally visualized via Schlieren-like, Mach and Pressure contours. It was observed that the usage of aero-disk spike reduced the aerodynamic drag by 45%–60% when compared to the bodies without spikes. The configuration labelled 3221 showed the maximum effect with an average drag reduction of more than 55%. There was a remarkable drop in the stagnation point value of surface heat flux with the addition of all five spike system. An efficient performance of localized peak heat fluxes was also noticed upon the addition of double and triple spike systems. In conclusion, it has been observed that the usage of double and triple spikes on the 3221 blunt body has significantly improved its aerodynamic drag and its surface heat flux.

The aim of this paper is to demonstrate the effects of the shape optimization on the missile performance at supersonic speeds. The N1G missile model shape variation, which decreased its aerodynamic drag and increased its aerodynamic lift at supersonic flow under determined constraints, was numerically investigated. Missile geometry was selected from a literature study for optimization in terms of aerodynamics. Missile aerodynamic coefficient prediction was performed to verify and compare with existing experimental results at supersonic Mach numbers using SST k-omega, realizable k-epsilon, and Spalart-Allmaras turbulence models. In the optimization process, the missile body and fin design parameters need to be estimated to design optimum missile geometry. Lift and drag coefficients were considered objective function. Input and output parameters were collected to obtain design points. Multiobjective Genetic Algorithm (MOGA) was used to optimize missile geometry. The front part of the body, the main body, and tailfins were improved to find an optimum missile model at supersonic speeds. The optimization results showed that a lift-to-drag coefficient ratio, which determines the performance of a missile, was improved about 11-17 percent at supersonic Mach numbers.

Aerodynamics occupies an important place in the design of high-speed ground transportation systems. When a vehicle is moving at a speed above 500 km/h under atmospheric pressure, the main energy is spent to overcome the aerodynamic drag. Creating a rarefied atmosphere inside a sealed pipe in order to significantly reduce energy loss is one of the key ideas of the HYPERLOOP project [1].The paper assesses the aerodynamic and thermal loads on the HYPERLOOP capsule fuselage in a partly evacuated tube based on the numerical solution of the Navier-Stokes equations of compressible flow closed by a differential turbulence model [2-4]. Numerical modeling was carried out with the help of the computational fluid dynamics software developed by the scientific researchers of the Institute of Transport Systems and Technologies of the National Academy of Sciences of Ukraine [5].It was shown that even under conditions of low air pressure in a partly evacuated tube the high-speed movement of the HYPERLOOP capsule will be accompanied by the formation of local supersonic zones, shock waves and non-stationary vortex systems. The structure of the flow essentially depends on geometry of the streamlined capsule and the speed of its movement.It was found that the flow structure and the values of aerodynamic dimensionless coefficients weakly depend on the pressure in the partly evacuated tube. Thus, the aerodynamic forces acting on the HYPERLOOP capsule at the same speeds are almost directly proportional to the pressure value in the tube.A certain problem in the design of the HYPERLOOP type high-speed vehicles will be the aerodynamic heating of the capsule fuselage. When the capsule moves at transonic speed the temperature of the outer surface of the capsule will be 60÷900 C. This heat load can have a negative impact on the performance of onboard power supply and control systems, as well as on the ensuring of the passengers’ comfort on the way.

Актуальность работы. Эффективность преобразования химической энергии топлива в механическое движение сверхзвукового летательного аппарата определяется потерями механической энергии на сопротивление летательного аппарата и потерями полного давления при торможении потока в воздухозаборнике воздушно-реактивного двигателя. Поэтому при прочих равных условиях более энергоэффективным будет тот летательный аппарат, у которого эти потери меньше. Существенное повышение энергоэффективности ожидается от применения осесимметричных изоэнтропических туннельных воздухозаборников. В рамках модели невязкого течения предлагается численная методика построения указанных воздухозаборников с профилированным центральным телом и цилиндрической обечайкой. Цель работы. В настоящее время отсутствуют способы профилирования осесимметричных изоэнтропических туннельных воздухозаборников, не имеющих элементов, выступающих за цилиндрический корпус летательного аппарата. Поэтому целью исследования является разработка методики построения таких воздухозаборников. Методы исследования. Используется численная реализация метода характеристик при условии изоэнтропичности течения. Предлагаемая методика состоит из двух задач. В первой задаче от угловой точки строится контур центрального тела с заданной ординатой точки фокусировки характеристик. Во второй задаче производится построение оставшейся части контура до другой угловой точки, при этом используется обращенное течение в кольцевом сопле с цилиндрической обечайкой. Одновременно с решением второй задачи определяется положение обечайки. Центральное тело может содержать на краях по угловой точке. При использовании промежуточной линии тока центральное тело будет гладким. Результаты. Создана методика для расчета семейства контуров осесимметричных туннельных сверхзвуковых воздухозаборников, геометрические характеристики которых однозначно описываются исходными данными.

In this paper, the flow control system consists of some small microflaps located between the rear fins of the projectile. These microflaps can alter the flow field in the finned region of the projectile resulting in asymmetric pressure distribution and thus producing control forces and moments, furthermore to provide directional control for a supersonic projectile. Due to the small size and high speed characteristics of projectile, which is with fast and valid response characteristics, this flow control system has initially shown excellent potential in terms of supersonic flow control. The CFD simulation used here solves steady-state Reynolds-averaged Navier-Stokes equation with two-equation turbulence model k-ε. Firstly, we investigate the flow mechanism around microflap in supersonic flow, the flow fields around the microflap are complex, involving three-dimensional shock-shock, shock-boundary layer interactions. Secondly, for the microflap and the fin of Basic Finner configuration, the influence of microflap geometric parameters, microflap locations on aerodynamics is obtained and the interference mechanism is explored. Finally, several typical roll and pitch control layouts are described. According to the simulation results and their analysis, some preliminary conclusions can be drawn: by analyzing the flow interference mechanism between microflap and the fin, we find that the separated shocks ahead of the microflap, the bow shocks around microflap, and the trailing-edge wake, have influences on fin's surface pressure; among these factors, the bow shocks are stronger than separated shocks, furthermore it can generate larger high pressure region. Then we find out the aerodynamic characteristics of several typical control layouts at a supersonic speed, Ma=2.5, furthermore, hence nearly 4.8% drag is increase compared with the condition without microflap. As the number of microflaps increasing, the control aerodynamic forces and moments increases almost linearly. With a proper layout of the microflap's location, quick change in the surface pressure distribution can be achieved for rear fins of the projectile, the microflap should be mounted that can increase the high pressure zone, meanwhile, reduce the low pressure zone on the surface of fins, thus modulating the projectile's attitude can be realized.

In this study, a computational investigation of multiphase flow within supersonic micronozzles has been performed in an effort to quantify the impact on thrust production and nozzle efficiency. Mot...

A body with blunt nose moving at supersonic speed generates a strong bow shock ahead of its nose. The bow shock creates high pressure on the blunt nose surface that increases aerodynamic drag on the body. The drag forcecan be substantially reduced if bow shock is replaced with oblique shockby attaching a spike in front of the blunt nose. Initially,in order to observe the effect of bluntness, various configurations are analyzed. It is observed that the drag force increases nonlinearly with bluntness. Next, a validation study is carried out to benchmark simulation methodology by reproducing published work through CFD. The purpose of present study is to optimize the shape and geometry of spike in order to reduce the drag force. A nose with small bluntness has no significant drag reduction with spike. Therefore a bluntness ratio of 0.7 was selected to see a significant effect of spike on drag reduction. A disk spike is selected from reference [1] that is validated with CFD. The effect of varying spike length on drag reduction is studied. A spike lengthis finalized that produces lower drag force at zero angle of attack over a supersonic Mach regime. Later on, the disk spike has also been analyzed with angle of attacks.

This manuscript presents a detailed characterization of active control of bow shock waves via leading edge injection, including subsonic coolant ejection and the appearance of Coanda effects. The flow phenomena occurring at steady and pulsating flow injection regimes were analyzed using steady and unsteady two-dimensional Reynolds-Averaged Navier Stokes, leading to a precise evaluation of the thermal load and drag reductions. Steady supersonic injection yields the largest abatement in thermal load and aerodynamic drag, while subsonic or fluctuating ones can also provide significant improvements at reduced cooling mass flow rates. Furthermore, a Coanda effect, causing a non-symmetric flow topology, was observed and analyzed for reduced injection port size. This Coanda effect is due to the sudden expansion happening from the injection port to the main flow and it causes the flow topology at the leading edge to become non-symmetric despite the complete symmetry of the problem. This is the first time in the literature such a phenomenon is documented for a supersonic airfoil leading edge injection. Furthermore, it enables the design of novel flow control strategies for the leading edge shock topology and flow structures in supersonic flows.

The flow structure and aerodynamic loads in evacuated tube transportation with three nose lengths (5, 10 and 15 m) at 1500 km/h are investigated by overset mesh technology. The governing equations of the IDDES model for 3-D, unsteady compressible flow are employed. The numerical algorithm is validated via moving model and wind tunnel experimental data. The flow topology considering the ground effect is revealed and compared. The wake is highly complex, characterized mainly by wave systems and trailing vortices. Additionally, the influence of the nose length on the generation of wave systems is discussed in detail. The largest shock intensity is identified at 5 m case, followed by 15 m case, and the smallest case at 10 m. Differences in wave systems also caused the two distinct mechanisms for the formation of flow separation and the evolution pattern of the wake vortex to be identified for three nose configurations. By comparison, the nose length increase is helpful to reduce the length of the disturbed region of the leading shock wave, the aerodynamic loads on the train surface and aerodynamic drag. A qualitative description is also presented of the relationship between the fluctuations of aerodynamic loads and flow structures.

Modern trends in the development of military and civil aviation are aimed at the constant growth of cruising and limit flight speeds of promising aircraft. The largest aviation powers (Russia, the USA, the EU, China, Japan) have created supersonic passenger and military aircraft and are talking about creating hypersonic aircraft. The pace of development of industrial technologies requires a significant reduction in the development of aviation equipment, so the advanced military aircraft of the 5th generation (Su-57, F-22, J-20, F-35) were developed within 10-15 years, and the requirements to the terms of creation of advanced aircraft complexes currently is no more than 8 years. These facts require a significant optimization of the aircraft life cycle stages, in particular, a significant volume of the development stage. In this paper, we consider the process of multi-criteria optimization of aircraft parameters at the stage of development and initial coordination of the minimum aerodynamic drag head-end of the aircraft fuselage, and assess the impact of this optimization on the life cycle of the product.

Computational Fluid Dynamics (CFD) analysis has been carried out on ballistic bodies with a hemispherical aero-disk spike. This research involves the presence and absence of a blunt tip to understand the flow behavior on these bodies under two scenarios namely zero-degree angle of attack and at supersonic speed (Mach 2.0). The aerodynamic drag and thermal heat transfer effects of these spikes on three different blunt-body configurations have been studied. The flow patterns on these configurations were three-dimensionally visualized through Schlieren images. It was observed that the usage of an aero-disk spike reduced the aerodynamic drag by 50%–60%. Additionally, peak surface heat flux reduced significantly. Similar drag reduction trends were noticed with the usage of the blunt tip version as well, but with further reduction in peak surface heat flux. In conclusion, the usage of blunt tip aero-disk spike on the configuration labeled 3221 has significantly improved the drag and surface heat flux effects.

Understanding the performance of penetrators and aerodynamic bodies of revolution (missiles, rockets, aircraft noses, etc.) requires a close look at the drag and the heat transfer characteristics at a wide range of supersonic flight conditions. This research utilizes computational study and compares the aerothermal loads of supersonic flows around a new penetrator geometry, derived based on the optimization of the nose factor, to those of other common projectile shapes: conical, tangent-ogive, and power series nose geometries. The abundance of research on 0.3-caliber projectile made the choice for this research simple in order to maximize our ability to compare to the existing data. The comparison of our 0.3 caliber cylindrical projectile with other geometries shows that within the range of 500–1500 m/s flight speed the new geometry has the lowest aerodynamic drag, lowest body temperature, and least amount of heating.

The supersonic (M∞ = 4.85) flow past a cylinder with a forward insertmade of a highlyporous cellular material is numerically modeled within the framework of the Reynolds-averaged Navier–Stokes equations. The air flow in the gas-permeable insert is described on the basis of a skeleton model of a highly-porous medium, whose determining parameters are the porosity coefficient (95%) and the pore dimensions (1 mm) of the actual cellular material. The aerodynamic drag coefficients of the model with different lengths of the porous forward insert are calculated on the unit Reynolds number range from 6.9 × 105 to 13.8 × 106 m−1. They are in agreement with the available experimental data, which indicates the adequacy of the proposed skeleton model in describing the actual properties of highly-porous materials.

The influence of a thermal wake due to gas injection and due to a pulsating optical discharge (POD) on the aero-dynamic-drag force of a body in a supersonic air flow with Mach number M = 1.45 are experimentally examined. With the help of a single-component aerodynamic balance, the influence of the injected subsonic jet and the thermal wake produced by POD on the aerodynamic drag of a hemisphere-on-cylinder model was studied. It is shown that the observed aerodynamic-force reduction phenomenon can be made more pronounced by increasing the laser power and pulse repetition frequency, or by decreasing the distance between the model and the pulsating optical discharge. The maximum aerodynamic-force reduction (up to 15%) due to the thermal-wake action was observed at a maximum laser-radiation power of W = 2.3 kW and at a pulse rate of f = 90 kHz. The joint effect due to the argon jet and due to the POD caused an aerodynamic-drag force reduction reaching 30%.

We have experimentally and numerically studied the influence of the ratio of the diameter of a cylinder with a frontal gas-permeable porous insert made of nickel sponge to the average pore diameter in the insert on the aerodynamic drag of this model body in supersonic airflow (M∞ = 4.85, 7, and 21). The analytical dependence of the normalized drag coefficient on a parameter involving the Mach number and the ratio of cylinder radius to average pore radius in the insert is obtained. It is suggested to use this parameter as a similarity criterion in the problem of supersonic airflow past a cylinder with a frontal high-porosity cellular insert.

Supersonic parachutes are a critical element of planetary mission whose simple structure, light-weight characteristics together with high ratio of aerodynamic drag makes them the most suitable aerodynamic decelerators. The use of parachute in supersonic flow produces complex shock/shock and wake/shock interaction giving rise to dynamic pressure oscillations. The study of supersonic parachute is difficult, because parachute has very flexible structure which makes obtaining experimental pressure data difficult. In this study, a supersonic wind tunnel test using two rigid bodies is done. The wind tunnel test was done at Mach number 3 by varying the distance between the front and rear objects, and the distance of a bundle point which divides suspension lines and a riser. The analysis of Schlieren movies revealed shock wave oscillation which was repetitive and had large pressure variation. The pressure variation differed in each case of change in distance between the front and rear objects, and the change in distance between riser and the rear object. The causes of pressure oscillation are: interaction of wake caused by front object with the shock wave, fundamental harmonic vibration of suspension lines, interference between shock waves, and the boundary layer of suspension lines.

Reducing the aerodynamic drag of bodies of revolution at supersonic speeds due to the creation of a bottom thrust is an important subject of aeromechanics. In the framework of this problem, it is proposed to perform numerical experiments on a model with a detachable bottom. It is assumed that the detachable bottom is affected by the main air flow jets separating from the surface of the body of revolution. The gas is reflected from the bottom and acts on the body of revolution, creating a bottom thrust and accelerating it.

Aerodynamic drag and heat reduction effectivity of the aerospike attached to the blunt-body at various aerospike semi-cone angles (θS), lateral injection from the aerospike stem, and a small bump on the aerospike stem, at different Mach number is numerically investigated. An open-source computational fluid dynamics code, i.e., rhoCentralFoam, a density-based solver in OpenFOAM is employed to solve the governing equations of supersonic turbulent flow. Menter's two-equation turbulence model, i.e., k−ω shear stress transport model is employed for turbulence modeling. A significant reduction in the total drag force (TDf) on the blunt-body is observed with the increase in aerospike θS at a fixed spike length (L)/blunt-body diameter (D) ratio for Mach 2 and 5. With the increase in θS>15° for L/D = 1 and θS>10° for L/D = 2, a significant decrease in the magnitude of coefficient of pressure is observed for Mach 5. Results show a maximum percentage reduction of 23.611% and 61.414% in TDf at L/D = 2 and θS=45° for Mach 2 and 5, respectively. Correlations are developed for the estimation of total drag force on the blunt-body and average surface temperature of the nose at Mach 2 and 5. Lateral injection substantially improves the aerodynamic heat reduction capability of the aerospike owing to the rapid expansion of the injectant in the main flow. An alternate passive technique (a small bump on the spike stem) capable of producing higher aerodynamic drag reduction compared to the active technique (i.e., lateral injection) is proposed. The small bump on the spike facilitates an early initiation of boundary layer separation and leads to the formation of a large recirculation zone ahead of the nose. Results indicate a higher reduction in aerodynamic drag with the increase in bump height (HB) compared to lateral and no injection at Mach 2 and 5. Present results have been validated with the experimental results available in the literature.

Drag and heat flux reduction is very important for the design of the hypersonic vehicle, and many novel strategies have been proposed recently. As one of the valid drag and heat flux reduction schemes, the forward-facing cavity has attracted an increasing attention for the scholars. In the current study, the influence of the cavity configuration on the drag and heat flux reduction mechanism of a blunt body has been investigated numerically by the two-dimensional axisymmetric Reynolds-averaged Navier–Stokes (RANS) equations coupled with the SST k–ω turbulence model, and the effects of the depth of the cavity and the angle of its trailing edge have been taken into consideration. At the same time, the numerical approaches employed in this paper have been validated against the available experimental data in the open literature, as well as the grid independency analysis. The obtained results shows that the cavity configuration has only a slight impact on the drag reduction of the blunt body in the supersonic flow with the freestream Mach number being 3.98, and the maximum drag reduction coefficient is within 1.0%. The forward-facing cavity with L / D = 4.0 and θ = 45 ° is the most beneficial configuration for the heat flux reduction of the blunt body in the range considered in the current study, and it should be validated by the wind tunnel test in the near future.

To date, in all the Mars landing missions, a supersonic parachute has been used to decelerate the capsule's entry into Mars' thin atmosphere from supersonic to subsonic speeds, because of its low mass and high aerodynamic drag. However, when the flexible porous parachute is placed in a supersonic flow, complex interdependent phenomena are observed around the two-body configuration system. These include aerodynamic interactions between the turbulent wake with shocks and coupled interactions between the compressible flows with the flexible porous structures undergoing large deformations. The flow instability around the parachute originates from these aerodynamic interactions between the canopy shock and the capsule wake, which depend on the Mach number, Reynolds number, proximity and size ratio of the canopy to the capsule, angles of attack of the canopy and capsule, and the material properties of the canopy. The effects of these parameters on the performances of the supersonic parachutes have been studied experimentally since the late 1950s. Computational fluid dynamics (CFD) and fluid-structure interaction (FSI) approaches have been used to understand the flow physics and the driving mechanisms of the aerodynamic interactions around the parachute system. However, the underlying mechanisms of these fully coupled interactions are not clearly understood. This review presents the recent developments in the study of supersonic parachutes based on the use of CFD and FSI simulations to predict their aerodynamic performances, summarizes the progress in experimental and numerical research aiming to investigate the unsteady aerodynamics in such a FSI problem, and concludes with discussions on the future challenges in the design and application of supersonic parachutes for Mars exploration missions.

Motivated by the growing scientific and engineering interest in evacuated tube railway transportation systems, in this paper we numerically study the influence of the vacuum degree on the flow field around a train capsule running in an evacuated tube with circular section. The vacuum degree is increased by lowering the nominal pressure inside the tube. The numerical simulations are fully verified by wind-tunnel experimental data of supersonic flows around a blunt body and in a scramjet combustion chamber, as well as by several numerical results in other related studies. The flow around the train capsule is characterized by a compression region in front of the train, a chocked flow near the train, and a complex highly unsteady region behind the train, where expansions waves and reflecting oblique shock waves exist. The total aerodynamic drag and the vacuum degree are found to be linearly related, revealing that lowering the nominal pressure can have a significantly beneficial effect on the aerodynamic performance of the train capsule. The aerodynamic heating due to compressibility effects and the increased pressure are more prominent along the centreline of the tube than on the tube wall. As the vacuum degree increases, the temperature and pressure differences between the front and the tail of the train and the intensity of the reflected shock waves become less significant, so that the extension of the expansion region in the train wake shortens.

A theoretical and experimental study has been undertaken, under ESA contract, by a team including GMV, the University of Bologna and EADS-Astrium, to assess the feasibility of an entry descent and landing (EDL) vehicle based on unpowered rotary decelerator. A deployable rotor with telescopic blades is considered, aimed at replacing all the standard deceleration devices for a descent module (parachutes, airbags, and retrorockets) except for the heat shield, with Mars as the main planetary target.This paper deals with the aerodynamic assessment of the proposed vehicle configuration. The objectives are twofold. First, a proof of concept of the main autorotation related events, viz. blades deployment from folded configuration in supersonic or transonic regime and telescopic blades extension in subsonic regime, is experimentally analyzed. Second, a characterization of the vehicle performances in terms of aerodynamic drag, lift to drag ratio and rotor angular speed is accomplished. Experiments are performed in the supersonic wind tunnel of Von Karman Institute of Fluid Dynamics and in the low speed wind tunnel of the University of Bologna. The vehicle is found to feature longitudinal static stability within the tested range of angles of attack. Multivalued rotor angular rates are found at subsonic speeds, corresponding to stalled and unstalled operation, with major differences in the achieved decelerating performances. Overall, the autorotation with telescopic blades is regarded as a viable concept for Mars EDL.

Aerodynamic drag, as a result of the high surface pressure generated because of the formation of a bow shock in front of a blunt nose cone of a hypersonic vehicle, is a serious problem. An aerospike can be attached to the stagnation point of the blunt nose cone to achieve a large reduction in drag and heat load at very high speeds in supersonic or hypersonic range. Aerospike replaces the strong detached shock wave with a system of weaker oblique shock waves, and a recirculation zone is formed between the shock and the body. Aerodynamic drag reduction on supersonic and hypersonic vehicles helps in reaching the desired range or altitude, allows economizing the fuel usage, simplifies the propulsion system requirements, and maximizes the ratio of payload to takeoff gross weight. In the present study, the flow around a blunted nose cone fitted with an aerospike is numerically simulated. An analysis is performed at a high Mach number (M=6) and different length-to-diameter ratios (L/D=1.5, 2) at different angles...

This paper describes the design optimisation study used to aerodynamically optimise the fairings that cover the rear wheels of the Land Speed Record vehicle, BLOODHOUND SuperSonic Car (SSC). Initially, using a Design of Experiments approach, a series of Computational Fluid Dynamics simulations were performed on a set of parametric geometries, with the goal of identifying a fairing geometry that was aerodynamically optimised for the target speed of 1,000 mph. Several aerodynamic properties were considered when deciding what design objectives the fairings would be optimised to achieve; chief amongst these was the minimisation of aerodynamic drag. A parallel, finite-volume Navier–Stokes solver was used on unstructured meshes in order to simulate the complex aerodynamic behaviour of the flow around the vehicle’s rear wheel structure, which involved a rotating wheel, and shockwaves generated close to a supersonic rolling ground plane. It was found that the simple response surface fitting approach did not sufficiently capture the complexities of the optimisation objective function across the high-dimensional design space. As a result, a Nelder–Mead optimisation approach was implemented, coupled with Radial Basis Function design space interpolation to find the final optimised fairing design. This paper presents the results of the optimisation study as well as indicating the likely impact this optimisation will have on the ultimate top speed of this unique vehicle.

Purpose This study aims to decrease the aerodynamic drag of the body of revolution at supersonic speeds. Supersonic area rule is widely used in modern supersonic aircraft design. Further reduction of the aerodynamic drag is possible in the framework of Euler and Reynolds averaged Navier–Stokes (RANS) equations. Sears–Haack body of revolution shape variation, which decreased its aerodynamic drag in compressible inviscid and viscous gas flow at Mach number of 1.8 under constraint of the volume with lower bound equal to volume of initial body, was numerically investigated. Design/methodology/approach Calculations were carried out in two-dimensional axisymmetric mode in the framework of Euler and RANS with SST model with compressibility correction equations at structured multiblock meshes. Variation of the radius as function of the longitudinal coordinate was given as a polynomial third-order spline through five uniformly distributed points. Varied parameters were increments of the radius of the body at points that defined spline. Drag coefficient was selected as an objective function. Parameter combinations corresponding to the objective function minimum under volume constraint were obtained by mixed-integer sequential quadratic programming at second-order polynomial response surface and IOSO algorithm. Findings Improving variations make front part of the body become slightly blunted, transfer part of volume from front part of the body to back part and generate significant back face. In the framework of RANS, the best variation decreases aerodynamic drag by approximately 20 per cent in comparison with Sears–Haack body. Practical implications The results can be applied for the aerodynamic design of the bullets and projectiles. The second important application is knowledge of the significance of the difference between linearized slender body theory optimization results and optimization results obtained by modern computational fluid dynamics (CFD) optimization techniques. Social implications Knowledge about the magnitude of the difference between linearized slender body theory optimization results and optimization results obtained by modern CFD optimization techniques can stimulate further research in related areas. Originality/value The optimization procedure and optimal shapes obtained in the present work are directly applicable to the design of small aerodynamic drag bodies.

Experimental and numerical studies were carried out to evaluate the effect of vortex generator on a small cylindrical protrusion at Mach number 2.0. The experiments were performed using the supersonic blow down wind tunnel on different heights of cylindrical protrusion with vortex generator placed ahead of them. The upstream and downstream flow around the cylindrical protrusion is influenced by vortex generator as is observed using both visualization and pressure measurement techniques. Numerical studies using three dimensional steady implicit formulations with standard k-ω turbulence model was performed. Results obtained through the present computation are compared with the experimental results at Mach 2.0. Good agreements between computation and experimental results have been achieved. The results indicate that the aerodynamic drag acting on cylindrical protrusion can be reduced by adopting vortex generator.

The reduction in aerodynamic drag by injecting a gaseous jet from the nose of a blunt body into a supersonic stream is investigated numerically. The penetration of the jet into the supersonic flow modifies the shock structure around the body and creates a low pressure recirculation zone, thereby decreasing the wave drag significantly. Combining various theoretical estimates of different flow features and numerical simulations, we identify a universal parameter, called the jet to freestream momentum ratio (RmA), which uniquely governs the drag on the blunt body. The momentum ratio fundamentally decides the penetration of the jet as well as the extent of low pressure envelope around the body. In addition, various influencing parameters reported in the literature are reviewed for different steady jet flow conditions. Furthermore, their limitations in regulating the flowfield are explained by correlating the facts with the jet to freestream momentum ratio. We perform the simulations for various combinations of physical and flow parameters of the jet and the freestream to show a universal dependence of drag on the momentum ratio.

Numerical analysis has been performed on an axisymmetric model with a spike at its tip. Steady simulation has been performed at supersonic speed of Mach 2. At high Mach regimes, aeroheating and high values of drag have always been a major concern for the aerospace industry. These two parameters have been subjected to decades of research, and as an outcome, design optimization has proven to be one of the most effective methods in reducing drag and heating levels. Revolving around the same issue, this paper analyzes various spike geometries on a blunt body and the drag coefficient values have shown the reduction of about 40–50% in the aerodynamic drag.

High speed aerodynamic bodies are associated with high aerodynamic drag. Reduction of this drag is a necessary requirement, for which different active and passive control techniques have been proposed. In the present work, the concept of breathing blunt nose (BBN) for drag reduction at hypersonic speed has been applied to a long blunt nosed body at supersonic speed. Investigation on two breathing configurations has been done using CFD tools and effect of breathing on static aerodynamic coefficients particularly on axial force coefficient has been studied. The employed breathing configurations increase normal force coefficient and pitching moment coefficient by 5% and 10% respectively. The axial force coefficient is reduced by 3.3% and 4.4% for the two breathing configurations at α=0°. The Ca reduction increases linearly with angle of attack, and it reaches 6.25% and 6.4% at α=8°. The results of non-breathing body (NBB) have been validated quantitatively using experimental data; and the results of breathing configurations match qualitatively with those found in the literature.

Passive and active flow control methods are used to manipulate flow fields to reduce acoustic signature, aerodynamic drag and heating experienced by blunt bodies flying at supersonic and hypersonic speeds. This paper investigate the use of active opposing jet concept in combination with geometric variations of the opposing jet nozzle to alleviate high wave drag formation. A numerical study is conducted to observe the effects of simple jet as well as jet emanating from a divergent nozzle located at the nose of a blunt hemispherical body. An initial discussion is presented of the complex shock wave pattern flow physics occurring when opposing jet ejected from a nozzle under various operating conditions interacts with the free stream flow. The complex flow physics that include long penetration and short penetration mode is studied in conjunction with effect on drag. The numerical setup consists of supersonic free stream flow interacting with an opposing sonic jet under varying pressure ratios. Initial computational results are validated by identifying prominent flow features as well as comparing available experimental data of surface pressure distributions. Preliminary validation is followed by the introduction of a divergent nozzle in the blunt body nose region. A series of numerical iterations are performed by varying nozzle geometric parameters that include nozzle divergent angle and nozzle length for a certain jet pressure ratio. Long penetration mode, short penetration mode as well as flow separations are captured accurately during the analysis. The results show a considerable reduction in drag by the use of a divergent nozzle. Specifically, 46% and 56% reduction in drag coefficient is achieved at pressure ratio of 0.6 and 0.8 respectively in the divergent nozzle cases as compared to the simple blunt body without any nozzle.

In this paper, the effects of boattail junction shaping on aerodynamic drag and stability of supersonic spin-stabilized rounds are investigated using computational fluid dynamics. For a generic round body comprising of a secant-ogive nose, a cylindrical body and a conical boattail, the shaping technique was achieved by adding a convex surface of varying degrees of radius of curvature to the junction between the cylindrical body and the boattail. It was shown through numerical simulations that this shaping technique can provide a reduction in aerodynamic drag of up to 5.4 % without destabilizing the round bodies when the radius of curvature is less than 8.8 times the diameter of the cylindrical body. The more gradual change of the flow characteristics, e.g., the pressure over the convex surface, was identified as the main reason for the drag reduction. A unique aspect of the current work is that stability is treated as an integral part of the performance assessment. It was also found that the dynamic instability encountered at large radii of curvature is due to the Magnus effects.

The objective of this is to analyse the flow field over an aerofoil section integrated with spikes at supersonic speed (Mach number greater than 1). Use of spike integrated with aerofoil changes the flow characteristics over aerofoil and hence aerodynamic lift and drag. The experiment consists of flow visualization graphs and measurement of coefficient of aerodynamic drag and lift. Here we are using different shapes of spike like sharp edge and hemi spherical edge. In this we will compare the flow over aerofoil with spike and without spike. The flow analysis is done by using Computational fluid dynamics (CFD). CFD is the study of external flow over a body or internal flow through the body. CFD is aiding aero-dynamist to better understand the flow physics and in turn to design efficient models. In short, CFD is playing a strong role as a design tool as well as a research tool.

Taking into account the issue of configuration or aerodynamic heating, most supersonic and hypersonic flight vehicles have to use the blunt-nosed body. However, in supersonic especially in hypersonic flow the strong bow shock ahead of the blunt nose introduces a rather high shock drag that affects the aerodynamic performance of the vehicles seriously. A spike mounted on a blunt body during its flight pushes the strong bow shock away from the body surface and forms recirculation flow with low pressure ahead of the body surface, and then decreases the drag. The drag reduction effects of spikes in high supersonic and hypersonic flow had been validated through experimental and numerical methods. In order to analyze the influence of the spike on aerodynamic characteristics at low supersonic (M=1.5) flow past blunt-nosed bodies, numerical studies were carried out which included the influence of the spike shape, the analysis of the fluid flow structures and the effect on the aerodynamic characteristics of a blunt body.

The article presents results of an experimental investigation of supersonic flow around of a longitudinal flow cylinder with a gas-permeable porous front insert. The main attention is paid to the influence of porous material characteristics of the front insert on the aerodynamic drag. The experiments were performed in wind tunnels at Mach number 4.85; 7 and 21. The used porous materials are a cellular-porous nickel, a fibrous porous material such as “metal felt” and a porous material formed from superimposed nets. The important role of the ratio of the cylinder diameter to the pore diameter in the aerodynamic drag of the model is shown for cellular-porous nickel. The data of numerical modeling of the flow around the model were used in order to reveal the physical mechanisms of the influence of the front porous insert on the aerodynamic drag.The article presents results of an experimental investigation of supersonic flow around of a longitudinal flow cylinder with a gas-permeable porous front insert. The main attention is paid to the influence of porous material characteristics of the front insert on the aerodynamic drag. The experiments were performed in wind tunnels at Mach number 4.85; 7 and 21. The used porous materials are a cellular-porous nickel, a fibrous porous material such as “metal felt” and a porous material formed from superimposed nets. The important role of the ratio of the cylinder diameter to the pore diameter in the aerodynamic drag of the model is shown for cellular-porous nickel. The data of numerical modeling of the flow around the model were used in order to reveal the physical mechanisms of the influence of the front porous insert on the aerodynamic drag.

The effect of an optical pulsed discharge created by CO2-laser with an average power of up to 2.2 kW on the aerodynamic drag of a model in a supersonic air flow is experimentally investigated. Experiments were carried out in a supersonic wind tunnel MAU-M (diameter of the nozzle outlet dc = 50 mm) on the modes M = 1.36, Re1 = 1.4– 3.8*107 1/m. To ensure a stable optical breakdown, a jet of argon gas was introduced into the focusing region of the laser beam. As a result, a decrease in the aerodynamic force acting on the model was obtained. This decrease was 15% when a subsonic flow of argon was introduced into the main flow and another 15% when optical pulsed discharge was organized.

The development of supersonic retropropulsion, an enabling technology for heavy payload exploration missions to Mars, is the focus of the present paper. A new experimental model, intended to provide computational fluid dynamics model validation data, was recently designed for the Langley Research Center Unitary Plan Wind Tunnel test section 2. Pretest analyses using modern computational fluid and thermal analysis tools were instrumental for sizing and refining the model, over the Mach number range of 2.4–4.6, such that tunnel blockage and internal flow separation issues would be minimized. A 5-in.-diam 70 deg sphere-cone forebody, which accommodates up to four 4∶1 area ratio nozzles, followed by a 9.55-in.-long cylindrical aftbody, was developed for this study based on the computational results. The model was designed to allow for a large number of surface pressure measurements on the forebody and aftbody. Supplemental data included high-speed schlieren video and internal pressures and temperatures. The r...

The effects of peripheral propulsive decelerator jets on the deceleration performance of Mars entry aeroshells are examined using computational fluid dynamics. The study considers a 70 deg blunt-cone aeroshell with four peripheral jets in Mach 12 flow of iodine-seeded nitrogen gas. Sonic and supersonic peripheral decelerator jets are considered in this study. The results show that the peripheral jets increase the standoff distance and change the profile of the bow shock around the aeroshell. The results also show that decelerator jets obstruct the flow around the aeroshell, which creates a low-pressure region on the surface behind the jets. The level of obstruction increases with thrust, which expands the low-pressure region. As a result, the aerodynamic drag coefficient of the aeroshell is inversely proportional to the thrust. The jet Mach number, however, has a small effect on the deceleration performance, with a maximum difference of less than 2% in the axial force between sonic and supersonic peripher...

Supersonic retropropulsion is an entry, descent, and landing technology applicable to and potentially enabling high-mass missions required for advanced robotic and human exploration on the surface of Mars. For conceptual design, it is necessary to understand the significance of retropropulsion configuration on an entry vehicle’s static aerodynamic characteristics and the relation of this configuration to other vehicle performance metrics. This investigation developed an approximate model for the supersonic retropropulsion flowfield to assist in evaluating the impact of design choices on the vehicle’s drag characteristics for flight-relevant conditions and scales. This model was used to explore the impact of operating conditions, required propulsion system performance, propulsion system composition, and vehicle configuration on the integrated aerodynamic drag characteristics of full-scale vehicles for Mars entry, descent, and landing. The forebody aerodynamic drag and axial force characteristics of vehicle...

An optimisation procedure coupled with computational fluid dynamics (CFD) is proposed to minimise the aerodynamic drag and to improve the static and dynamic stabilities of generic rounds at supersonic speeds (Mach 1·5 to 4). First, the Active-set algorithm, Sequential Quadratic Programming (SQP) is used as the optimisation method for drag minimisation. The objective function is the zero-lift drag computed from a semi-empirical solution. The constraints are based on the geometric restrictions of the body. CFD is then employed to validate the accuracy of the drag prediction from the semi-empirical solution and to incorporate the stability requirements into the optimisation process. A supersonic round body is considered as an example application. The optimised body provides up to 15% drag reduction and 46% increase in gyroscopic stability while remaining dynamically stable over the whole range of the operating Mach numbers.

A three-dimensional numerical simulation of a supersonic free-stream at Mach 2.5 over a spherical body with a sonic opposing jet from its stagnation point is carried out by solving the three-dimensional Navier–Stokes equations coupled with the standard k–ɛ turbulence model. It is aimed to investigate the effects of the jet on the drag reduction on the body and the flow field around the body. The influences of the jet pressure, the nozzle size of the jet, and the angle of attack are systematically studied for the purpose. An unsteady oscillatory motion mode and a nearly steady motion mode are identified depending upon the jet total pressure. There exists a critical jet pressure where the flow mode transition from one to the other happens suddenly and this critical pressure value varies approximately linearly with the jet nozzle exit size inversely. For the zero angle of attack, the results show that there exists a maximum overall drag reduction as the jet pressure changes for each jet nozzle size and the maximum overall drag reduction always happens at the unsteady oscillatory motion mode. The main shock in front of the body is pushed backward by the jet and the displacement of the shock decreases with the increase of the angle of attack, and the drag reduction efficiency also decreases with the angle. Regarding to the mode transition, it is found that the drag rises suddenly when the transition happens for the angle of attack smaller than or equal to 5° but it does not result in the rise for the angle larger than 5°. The results show that the maximum overall drag reduction can be reached as high as 32.6% for the cases studied. The present results provide useful information for drag reduction applications using an opposing jet.

In this study, a transonic flow past NACA0012 profile at angle of attack α=00 whose aspect ratio AR is 1.0 with non-equilibrium condensation is analyzed by numerical analysis using a TVD scheme and is investigated using an intermittent indraft type supersonic wind tunnel. Transonic flows of 0.78–0.90 in free stream Mach number with the variations of the stagnation relative humidity(Φ0) are tested. For the same free stream Mach number, the increase in Φ0 causes decrease in the drag coefficient of profile which is composed of the drag components of form, viscous and wave. In the case of the same M∞ and T0, for more than Φ0=30%, despite the irreversibility of process in non-equilibrium condensation, the drag by shock wave decreases considerably with the increase of Φ0. On the other hand, it shows that the effect of condensation on the drag coefficients of form and viscous is negligible. As an example, the decreasing rate in the drag coefficient of profile caused by the influence of non-equilibrium condensation for the case of M∞=0.9 and Φ0 =50% amounts to 34%. Also, it were turned out that the size of supersonic bubble (that is, the maximum height of supersonic zone) and the deviation of pressure coefficient from the value for M=1 decrease with the increase of Φ0 for the same M∞.

A matrix of simulations of hypersonic flow over blunt entry vehicles with steady and pulsing retropropulsion jets is presented. Retropropulsion in the supersonic domain is primarily designed to reduce vehicle velocity directly with thrust. Retropropulsion in the hypersonic domain may enable significant pressure recovery through unsteady, oblique shocks while providing a buffer of reactant gases with relatively low total temperature. Improved pressure recovery, a function of Mach number squared and oblique shock angle, could potentially serve to increase aerodynamic drag in this domain. Pulsing jets are studied to include an additional degree of freedom to search for resonances in an already unsteady flow domain with an objective to maximize the time-averaged drag coefficient. In this paradigm, small jets with minimal footprints of the nozzle exit on the vehicle forebody may be capable of delivering the requisite perturbations to the flow. Simulations are executed assuming inviscid, symmetric flow of a perfect gas to enable a rapid assessment of the parameter space (nozzle geometry, plenum conditions, jet pulse frequency). The pulsed-jet configuration produces moderately larger drag than the constant jet configuration but smaller drag than the jet-off case in this preliminary examination of a single design point. The fundamentals of a new algorithm for this challenging application with time dependent, interacting discontinuities using the feature detection capabilities of Walsh functions are introduced.