Cylindrical Afterbody Research Articles

Numerical Study of Supersonic Flow Past a Cylindrical Afterbody

The near wake of a cylindrical afterbody aligned with a uniform Mach 2.46 flow has been numerically investigated using Reynolds-Averaged Navier-Stokes equations (k-epsilon two equation model) and Large Eddy Simulation (dynamic sub-grid scale eddy viscosity model). Mean flow field properties obtained from numerical simulations, such as axial velocity, pressure on the base surface has been compared with the experimental results. It has been found that k-epsilon model fails to predict the flow properties in the recirculation region where better agreement has been observed between the data obtained from LES and measurements. Data obtained from LES has been further analyzed to investigate the turbulent flow field in the wake region. Parameters like turbulent kinetic energy and primary Reynolds stress have been calculated and compared with the results obtained from an experiment in order to achieve a better understanding of the role of turbulence in the flow field.

Investigation of the effect of nose shape and geometry at supersonic speeds for missile performance optimisation

This paper investigated and analysed the influence of the missile nose shapes and geometry on the drag force coefficients at supersonic Mach numbers ranging from 2 to 5. The nose shape investigation includes a pointed cone, pointed ogive, and blunt cone with a constant length and diameter of cylindrical afterbody. Subsequently, the nose length was manipulated to investigate how the nose geometry with different fineness ratios affected each drag component. The drag force coefficient and its components were determined using a program called MATLAB_Aerody, in which a semi-empirical approach was implemented for the calculations. The program was validated against the semi-empirical results with existing theoretical and experimental results. The comparison showed a maximum error of 3.6%. The results of the nose shape analysis show that a sharper and more slender nose with a pointy tip produces a minimal drag. Nose bluntness, however, creates considerable additional drag force due to a higher wave drag coefficient by forming detached or bow shock waves. At supersonic speeds, increasing the fineness ratio significantly reduces the drag force by drastically decreasing the wave drag coefficient with a marginal increase in the skin friction coefficients. However, the results show an insignificant additional gain for the skin friction drag and a reduction in the wave drag coefficient when the nose has a fineness ratio beyond 4.5. Furthermore, the results from both analyses show that the base drag coefficient is invariant with the nose shape and the fineness ratio. Therefore, a better missile performance can be achieved with a sharper and more slender nose shape with a higher fineness ratio.

Numerical investigation of a near-wake flowfield with base bleed

A computational study of a supersonic flow in the base region and the nearest wake of a cylindrical body moving at a supersonic speed has been carried out. In this case, the “prehistory” of the flow was taken into account, i.e. the configuration of the computational domain was as close as possible to the real one. The use of various RANS turbulence models for calculating flow in the base region and the nearest wake was analyzed. The following turbulence models were considered: 1) SST model; 2) Standard k–ε model; 3) k–ε RNG model; 4) Standard Reynolds Stress (RS) model; 5) RS BSL model. Based on a comparison of the calculation results with experimental data, it is shown that: 1) when calculating the flow in the base region and in the wake of the vehicle, it is very important to take into account the “prehistory” of the flow, i.e. to calculate the flow around the entire vehicle; 2) the best match was obtained using the the k–ε RNG model. The effect of base bleed on the near-wake flowfield of a cylindrical afterbody aligned with a Mach 2.5 flow has been investigated.

Numerical Investigation of a Supersonic Flow in the Near Wake Region of a Cylindrical Afterbody

A computational study of a supersonic flow in the base region and the nearest wake of a cylindrical body moving at a supersonic speed have been carried out. A mathematical model of high-enthalpy flows is presented. In this case, the "prehistory" of the flow was taken into account, i.e., the configuration of the computational domain was as close as possible to the real one. The use of various turbulence models for calculating flow in the base region and the nearest wake was analyzed. The following turbulence models were considered: 1) the Spalart --- Allmaras model; 2) SST model; 3) standard k--ε model; 4) k--ε model with compressibility correction; 5) k--ε RNG (renormalized group) model; 6) k--ε Realizable model; 7) standard Reynolds Stress (RS) model; 8) RS BSL (Reynolds stress baseline) model. Based on a comparison of the calculation results with experimental data, it is shown that: 1) when calculating the flow in the base region and in the wake of the vehicle, it is very important to take into account the "prehistory" of the flow, i.e., to calculate the flow around the entire vehicle; 2) the best match was obtained using Reynolds Stress models and the k--ε RNG model

Mitigation of Antisymmetric Mode and Drag With Base Cavities

Abstract Experiments were carried out to examine the impact of base cavities on the base pressure fluctuations and total drag of a cylindrical afterbody for freestream Mach numbers 0.6–1.5. Significant improvement in the base pressure and a substantial reduction in the afterbody drag was noticed in the presence of a base cavity at subsonic Mach numbers. However, on increasing the cavity length beyond a certain value, its performance deteriorates. At supersonic Mach numbers, their effectiveness drops drastically. Tones in the spectra can be classified into two types depending on the dominant azimuthal mode, which is either 0 or 1 and are referred to as symmetric and an antisymmetric mode, respectively. Spectra at subsonic Mach numbers exhibit tones, which are related either to mode 0 or 1. However, at supersonic Mach numbers, only tones related to mode 0 exist. The base cavity either effectively suppresses the antisymmetric mode or modifies it into a symmetric mode resulting in mitigation of the tones related to antisymmetric mode.

On Unsteady Flow Analysis of a Round Spike Blunt Nose Afterbody in Mach 6 Flow

An aerospike, in front of a blunt body, has largely been deemed as a key passive control device for effectively reducing the wave drag and aerodynamic heating associated with high-speed flows. In addition, it has been reported that the presence of a spike brings in unsteadiness in the form of oscillation and pulsation to the flow characteristics. Past researchers, having investigated mainly the aerothermodynamic coefficients, have hinted towards the suppressing of such oscillations with the use of a round nose spike over a sharp spike, though a thorough and a concrete result is yet to be established together with offering a detailed explanation of the flow physics and its dependence on the spike’s geometric characteristics (spike-length to afterbody-diameter ratio, L/D). Numerical investigation has been carried out using axisymmetric Navier-Stokes laminar flow solver at hypersonic Mach number of 6.0. A round-tip spike with a flat face cylindrical afterbody have been simulated for spike length ratios of L/D = 0.5 – 2.0, with spike diameter to-afterbody diameter (d/D) of 0.1. The behavior and subsequent control of flow pulsation for a round (hemispherical) spike with varying L/D-ratio has been established.

Effects of Forebody Geometry on Side Forces on a Cylindrical Afterbody at High Angles of Attack

The aim of the present study was to investigate the effects of the forebody geometry on the asymmetric side forces, separations and vortex shedding behind the vehicles at high angles of attack. It reports wall pressure measurements to describe the asymmetric vortex system developing over an inclined cylindrical body exposed to a low-speed freestream.

Modes of Base Pressure Fluctuations: Shape, Nature, and Origin

An experimental study was carried out to examine the unsteady pressure field on the base of a cylindrical afterbody at freestream Mach numbers ranging from high subsonic to low supersonic speeds. The objective was to gain insights into spatial modes of base pressure fluctuations by identifying their shape and nature, and the mechanism of their origin using spectral proper orthogonal decomposition. It is observed that 90–95% of the cumulative energy content resides within the first three modes at all Strouhal numbers for all freestream Mach numbers. Further, at subsonic Mach numbers and for Strouhal numbers , the shape of the modes indicates that their nature is symmetric, and the mechanism for the origin of the dominant mode is the pulsing action of the recirculation bubble. However, for Strouhal numbers , the shape of the modes is asymmetric, indicating that their nature is antisymmetric, and the mechanism associated with the dominant mode is the radial displacement of the recirculation bubble coupled with the motion of structures within the recirculation bubble. At supersonic Mach numbers and for all Strouhal numbers, the shape of the modes indicates that they are symmetric, and the dominant mechanism is the pulsing action of the recirculation bubble.

Mach Number Effect on Symmetric and Antisymmetric Modes of Base Pressure Fluctuations

An experimental study aimed at evaluating the influence of Mach number on the base pressure fluctuations of a cylindrical afterbody was performed over a wide range of Mach numbers from subsonic to supersonic speeds. Time-averaged results indicate that the coefficient of base pressure drops with the increase in the freestream Mach number at subsonic speeds and increases at supersonic Mach numbers. The coefficient of root-mean-square of the pressure fluctuations follows a decreasing trend with the increase in the Mach number. Examination of the spectra reveals different mechanisms dominate the pressure fluctuations from the center to the periphery of the base as well as with the change in the Mach number. Analysis of the azimuthal coherence indicates that all the dominant tones in the spectra can be classified either into a symmetric or an antisymmetric mode at subsonic Mach numbers. However, at supersonic Mach numbers, all the dominant tones in the spectra are symmetric in nature. The results from the cross-correlation suggest that two possible mechanisms of recirculation bubble pulsing and convective motions/vortex shedding are driving the dynamics on the base at subsonic Mach numbers. However, at supersonic Mach numbers, only single mechanism of the recirculation bubble pulsing dominates. Moreover, it indicates that the symmetric mode is associated with the dynamics of the recirculation bubble and the antisymmetric mode is related to the convective motions/vortex shedding.

Fluid-structure interaction of a rolling restrained body of revolution at high angles of attack

The current work investigates numerically rolling instabilities of a free-to-roll slender rigid-body of revolution placed in a wind tunnel at a high angle of attack. The resistance to the roll moment is represented by a linear torsion spring and equivalent linear damping representing friction in the bearings of a simulated wind tunnel model. The body is subjected to a three-dimensional, compressible, laminar flow. The full Navier-Stokes equations are solved using the second-order implicit finite difference Beam-Warming scheme, adapted to a curvilinear coordinate system, whereas the coupled structural second order equation of motion for roll is solved by a fourth-order Runge-Kutta method. The body consists of a 3.5-diameter tangent ogive forebody with a 7.0-diameter long cylindrical afterbody extending aft of the nose-body junction to x/D = 10.5. We describe in detail the investigation of three angles of attack 20°, 40°, and 65°, at a Reynolds number of 30 000 (based on body diameter) and a Mach number of 0.2. Three distinct configurations are investigated as follows: a fixed body, a free-to-roll body with a weak torsion spring, and a free-to-roll body with a strong torsion spring. For each angle of attack the free-to-roll configuration portrays a distinct and different behavior pattern, including bi-stable limit-cycle oscillations. The bifurcation structure incorporates both large and small amplitude periodic roll oscillations where the latter lose their periodicity with increasing stiffness of the restraining spring culminating with distinct quasiperiodic oscillations. We note that removal of an applied upstream disturbance for a restrained body does not change the magnitude or complexity of the oscillations or of the flow patterns along the body. Depending on structure characteristics and flow conditions even a small rolling moment coefficient at the relatively low angle of attack of 20° may lead to large amplitude resonant roll oscillations.

Numerical investigation of flow structures around a cylindrical afterbody under supersonic condition

Large Eddy Simulation (LES) with dynamic Smagorinsky model has been applied to numerically investigate the complicated flow structures that evolve in the near wake of a cylindrical after body aligned with a uniform Mach 2.46 flow. Mean flow field properties obtained from numerical simulations, such as axial velocity, pressure on base surface, have been compared with the experimental measurements as well as with other published results. It has been found that standard k–epsilon model fails to predict the flow properties in the recirculation region where better agreement has been observed between the data obtained from LES and experimental measurements. Flow Statistics like turbulent kinetic energy and primary Reynolds' stress have also been calculated and compared with the results obtained from experiments in order to quantitatively assess the ability of LES technique to predict the turbulence properties of flow field in the highly compressible shear layer region. The data obtained from LES has been further analyzed to understand the evolution of coherent structures in the flow field. Proper Orthogonal Decomposition (POD) of the data obtained from central plane in the wake region has been performed in order to reveal the most energetic structures present in the flow field.

The prediction of axial aerodynamic coefficient reduction using base bleed

To reduce axial aerodynamic force as leading aerodynamic load during projectile flight, one of the possible solutions is to redesign the projectile base, using limited flow as base bleed. Base bleed design needs accurate turbulent supersonic base flow model over a cylindrical afterbody and needs to be developed. Predictions and analysis of the aerodynamic behavior of the base bleed projectile with such a model can reduce development cost and test effort. There are three ways of achieving results of a suitable flow model: by theoretical prediction, by numerical simulation and by restricted experimental measurement. The base bleed effects on the aerodynamic axial coefficient, as increase of the base pressure, and decrease of the eddy component of axial aerodynamic coefficient. Results of the axial aerodynamic coefficient of projectile, without base bleed obtained by experimental and predictional calculation. A validation calculation was performed for several types of projectiles. Computed results show reasonable trends in the base pressure increase – base drag reduction, base corner expansion and downstream wake closure location.

Reynolds-Averaged Navier-Stokes/Large-Eddy Simulations of Supersonic Base Flow

Several zonal and nonzonal hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation approaches have been assessed to handle a high Reynolds number supersonic base flow. The results obtained on 5 and 13.5 x 10 6 points grids are compared to the available experimental data (Herrin, J. L., and Dutton, J. C., Supersonic Base Flow Experiments in the Near Wake of a Cylindrical Afterbody,AIAA Journal, Vol.32, No. 77,1994), and the capabilities of these different methodologies to predict supersonic flows are discussed. The highly compressible separated shear layer proved to be a challenging issue for hybrid methods due to an alteration in the instability process (as compared to the incompressible case), leading to three-dimensional coherent structures. Numerous numerical parameters relevant to hybrid methods have been assessed. The incoming boundary layer thickness needs to be properly set, while a weak influence of the subgrid scale model is observed when small-scale structures in the separating mixing layer are resolved. Another finding of the present study is the dramatic influence displayed by the numerical dissipation on the flowfield. Sensitivity to the C DES model constant is observed even with the finest grid, leading to a delay in the generation of instabilities in the shear layer.

Microactuators for Projectile Flight Control Systems: a Feasibility Study

New concepts for microactivated flight control of projectiles have been investigated for operation in both subsonic and supersonic flight regimes. A successful concept can provide an inexpensive flight control device with many control redundancies, and it can reduce size weight and energy consumption of the control system. Two promising concepts have been explored in detail. The first one consists of a microflap in the form of a delta wing of 3 mm base by 4 mm height, which can be deployed and activated at a location on the ogive of the projectile or at the beginning of its cylindrical afterbody. The second concept is based on microballoons, which can be inflated and change the local pressure distribution on the projectile. A computational-fluid-dynamics (CFD) computer code has been used to model these two flight control concepts

Wake Closure and Afterbody Heating on a Mars Sample Return Orbiter

Aeroheating wind-tunnel tests were conducted on a 0.028-scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near-wakeclosure and to determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley Research Center 20-Inch Mach 6 Air and CF 4 Tunnels for postnormal shock Reynolds numbers (based on forebody diameter) ranging from 1.4 × 10 3 to 4.15 x 105, angles of attack ranging from -5 to 10 deg at 0-, 3-, and 6-deg sideslip, and normal shock density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75 % of the reference forebody stagnation point heating.

Velocity Measurements in a Pressure-Driven Three-Dimensional Compressible Turbulent Boundary Layer

The flow characteristics of a three-dimensional, compressible, turbulent boundary layer have been investigated experimentally. The three dimensionality was generated by inclining a cylindrical afterbody at 10-deg angle of attack to a Mach 2.45 freestream. The objective was to determine the mechanisms that govern the growth and behavior of pressure-driven, three-dimensional, compressible, turbulent boundary layers. Laser Doppler velocimetry was used to determine mean velocity components and turbulence statistics

Two-Equation k-s Turbulence Model: Application to a Supersonic Base Flow

Base e ows are an important subject of theoretical studies due to the determinant role of the rear part of projectiles, missiles, or launchers in their e ight capabilities. The problem of accurately predicting the aerothermal e elds in which they are embedded is still not satisfactorily solved by current theoretical methods. However, such e ows are very useful test cases for improving turbulence modeling because they incorporate nearly all of the fundamental problems that can be encountered in supersonic turbulent e ows submitted to strong viscous interactions. Axisymmetriccone gurationsareappreciableforperforming numerousparametricalstudiesofmodelsdueto their rigorous two dimensionality. In this framework, a theoretical examination of the k‐! model is performed, and its principal distinguishing characteristics with respect to k‐ models are used to establish a k‐ae model whosesecond variable has a simple physical interpretation. A closeexamination of the predictivecapabilities of thesethree types of models is done in comparison with the well documented and largely accepted database from the experiments of Herrin and Dutton (Herrin, J. L., and Dutton, J. C., “ Supersonic Base Flow Experiments in the Near Wake of a Cylindrical Afterbody,” AIAA Paper 93-2924, July 1993 ). In particular, the evolution of the e uctuating quantities predicted by the models in separated e ow are given in detail. The results obtained prove that the simulation of separated e ows with two-equation turbulence models is still promising in the framework of Reynolds averaged Navier‐Stokes calculations.

Velocity and turbulence measurements in a supersonic base flow with mass bleed

Two-component laser Doppler velocimetry was used to obtain detailed mean velocity and turbulence measurements in the near wake of a cylindrical afterbody with base bleed in a Mach 2.5 flow. The bleed flow provides at least some of the fluid required for shear layer entrainment and shields the base annulus from the outer shear layer and the primary recirculation region, leading to an increase in base pressure. There is an overall reduction in turbulence levels throughout the base bleed flowfields relative to the near-wake flowfields of blunt-based and boattailed afterbodies. With increasing bleed, the formation of a strong bleed jet shear layer and secondary recirculation region near the base annulus offsets the benefits of base bleed, leading to a drop in the base pressure. The net benefits of base bleed are maximized at the optimum bleed condition, which corresponds to the highest base pressure, the disappearance of the primary recirculation region, and the lowest turbulence levels in the near-wake flowfleld. Increased benefits from base bleed could be achieved by injecting the bleed fluid at the lowest possible velocity through the use of larger bleed orifices, porous bases, or bleed orifices located along the outer base annulus.

Similarity rule for jet-temperature effects on transonic base pressure

On the basis of the similarity rule for jet interaction, a hot jet can be simulated in a cryogenic wind tunnel with a test gas at ambient or moderately elevated temperatures. By using this approach, jet-temperature effects on the base pressure of a cylindrical afterbody model at transonic speeds have been investigated in the 0.1-m transonic cryogenic wind tunnel at the National Aerospace Laboratory. Mixtures of nitrogen and either methane, argon, or helium at varying temperatures were used as a jet gas to determine separate effects of jet temperature, specific heat ratio, and gas constant. It has been found that data obtained for various jet conditions can be correlated very well with two similarity parameters : the plume maximum diameter for the plume shape effect and the jet to freestream mass flux ratio for the jet entrainment effect. To verify this similarity rule, the same model was tested in an ambient wind tunnel. It was found that an ambient temperature gas having low molecular weight could simulate the jet temperature effects on the transonic base pressure.

Supersonic near-wake afterbody boattailing effects on axisymmetric bodies

An experimental investigation of the near-wake flowfield downstream of a conical boattailed afterbody in supersonic flow is presented. The afterbody investigated is typical of those for conventional boattailed missiles and projectiles in unpowered flight. Flow visualization, mean static pressure measurements, and three-component laser Doppler velocimeter data have been obtained throughout the near wake of the body. The effects of afterbody boattailing on the physics of the near-wake flow are determined by comparing the present data with similar data obtained on a cylindrical afterbody. Results indicate that a net afterbody drag reduction of 21 % is achieved with the current boattailed afterbody for a freestream Mach number of 2.46. The shear-layer growth rate, and therefore mass entrainment from the recirculation region behind the base, is shown to be significantly reduced by afterbody boattailing due to the reduction in turbulence levels throughout the near wake as compared to the cylindrical afterbody.