Drag reduction for payload fairing of satellite launch vehicle with aerospike in transonic and low supersonic speeds

Time-dependent turbulent compressible Reynolds-averaged Navier–Stokes equations are solved for computing flowfield over the bulbous payload shroud of a satellite launch vehicle at a freestream Mach number of 0.95 and at a zero angle of attack. Numerical simulation has been carried out employing a multistage Runge–Kutta time-stepping scheme in conjunction with a finite volume discretization. The closure of these equations is achieved using the Baldwin–Lomax turbulence model. Comparisons have been made between numerical and experimental results such as schlieren pictures, position of terminal shock on forebody cylinder and surface pressure distributions. A good agreement is found between them. A separated flow zone on the boat tail region is observed and is found to be having unsteady pressure fluctuations. Standard deviation, higher-order moments and spectrum of surface pressure levels of the fluctuating surface pressure are analysed in the separated flow region of the boattail of the heat shield. A comparative study of surface pressure fluctuations and spectrum of sound pressure levels has been made for two different types of heat shield configurations that are close to the recommended NASA specifications.

The article presents a flow field visualization study on a typical payload shroud at transonic Mach numbers range. The numerical simulation over the heat shield is carried out by solving axisymmetric, turbulent, compressible, Reynolds-averaged Navier—Stokes equations. The closure of these equations is achieved using the Baldwin—Lomax turbulence model. The density contour plots are obtained at Mach number 0.80, 0.90, 0.95, and 1.0. The transonic flow field features over the heat shield are captured and compared with the schlieren pictures, and good matching between them is exhibited. The movement of the terminal shock is found to be a non-linear function of freestream Mach number. The location of the shock on the heat shield is also confirmed with the flight data. The numerical studies have shown that the reliable estimation of the transonic flow field can be made in conjunction with the flow visualization technique.

Single-mode flutter is a type of panel flutter occurring at transonic and low supersonic speeds. Transition to instability in the form of single-mode flutter occurs without interaction between natural modes, in contrast to coupled-mode flutter, where coupling between the 1st and 2nd eigenmodes takes place. Blade flutter is one of the main issues that engine designers have to face. The danger of this phenomenon is a rapid increase of blades stresses, which can lead to their destruction. In this paper, a single-mode flutter of panels of rectangle and parallelogram panels and the influence of various design parameters on blade flutter boundary are investigated with using the energy method. It is shown that for parallelogram plates even at a small skew angle the aeroelastic stability increases significantly at transonic and low supersonic flight speeds. The study of blade flutter showed that effect of the inter-blade tension in the mid-span shroud on the flutter is significant in contrast to other investigated parameters.

The main focus of the present paper is to computational fluid dynamics analysis and design of payload fairing of satellite launch vehicle at freestream Mach number range of 0.6 - 3.0. Initially, time-dependent compressible three-dimensional Euler equations are solved employing a finite volume discretization method with a multi-stage Runge-Kutta time-stepping scheme to compute surface pressure and aerodynamic coefficients at various payload fairing and at angle of attack up to 5o with an increment of 1o. Payload fairing dimensions are selected that satisfies permissible structure load on satellite launch vehicle. Detailed flowfield simulation is carried out on the selected payload fairing employing axisymmetric compressible Reynolds-average Navier-Stokes equations to assess unsteady flowfield characteristics. The numerical simulations are used to locate terminal shock on the payload fairing at transonic Mach number. Unsteady flow characteristics are used to compute acoustic load. Shock standoff distances at supersonic speeds are tabulated and compared with the analytical solution. Schlieren images and oil flow pictures are compared with experimental results and in good agreement. Aerodynamic shape optimization of satellite launch vehicle payload fairing shape has been performed to satisfy structural load at maximum drag and dynamic pressure.

Comparative study of surface pressure fluctuations over bulbous heat shields at Mach number=0.95

Solution of viscous flow over heat shield using structured grid

An aerospike projectile base is an inverted rocket nozzle shape intended to reduce the base drag on a projectile by forcing isentropic flow over a portion of the projectile base and reducing the size of the base recirculation region. Initial studies on aerospike bases showed a significant reduction in overall drag compared to a control projectile shape with the same ogive shape. Both the aerospike and control projectile were fabricated as monolithic copper based upon the shape of a well-studied Sierra 168 grain Boat Tailed Hollow Point (BTHP) reference projectile. It was observed in testing that the overall aerodynamic drag of the monolithic control projectile was significantly higher than the published drag of the reference projectile despite being ballistically similar with the exception of the hollow ogive cavity (hollowpoint). The original aerospike study was recreated with the addition of drilled hollow points in the control and aerospike projectiles. A statistically significant reduction of approximately 4% in aerodynamic drag was measured in both the control and aerospike projectile at Mach numbers greater than 2 in testing. No reduction in drag was observed in Mach numbers below 2. The hollowpoint did stabilize the control projectile at low supersonic and transonic speed allowing a better comparison of the aerospike to the control projectile.

Aerodynamic shape optimization of launch vehicle hammerhead payload fairing (PLF) boat tail shapes has been carried out. Multi-objective optimization studies have been carried out to minimize the boat tail length, drag coefficient, the length of separated flow behind the boat tail, and the maximum pressure ratio over the boat tail. The boat tail shape has been parameterized using a B-spline. A novel optimal shape has been obtained, which has a shallow initial slope followed by a rapid increase in the slope near the boat tail end, termed the ramp stepped boat tail (RSBT). The optimization study has been carried out by varying the stage-to-PLF diameter ratios. A novel parameterization strategy has been employed, which convexifies the design space. Computational fluid dynamics (CFD) simulations have been carried out on 950 different boat tail configurations sampled using Latin hypercube sampling (LHS) at Mach number 0.8. Results from CFD have been validated against experimental data from literature. Artificial neural networks have been used as surrogate models for computing the objectives. Multi-objective optimization was carried out using the genetic algorithm for various stage-to-PLF diameter ratios. Pareto-optimal fronts obtained have been analyzed, and optimal shapes derived from them have been validated using CFD. It is found that the optimal RSBT configuration is found to be advantageous over the conventional conical boat tail in reducing drag, separated flow length, boat tail length, shock strength, and the turbulent kinetic energy downstream of the boat tail.

A numerical study is made to analyze the performance of a secant-ogive-cylinder projectile in the transonic regime in terms of aerodynamic drag. At transonic speeds, the base drag contributes a major portion of the total aerodynamic drag, and hence affects projectile's performances significantly. The base bleed method is applied to reduce the base drag by varying the value of parameters, the bleed quantity (I) and the bleed area ratio (ϖ). The implicit, diagonalized, symmetric Total Variation Diminishing (TVD) scheme, accompanied by a suitable grid, is employed to solve the thin-layer axisymmetric Navier-Stokes equations coupled with the Baldwin-Lomax turbulence model. The computed results show that, in comparison with the case without base bleed, an increase in bleed quantity or a higher injection speed due to a smaller bleed area ratio at fixed bleed quantity can result in a base (and total) drag reduction. At Mach number 0.96, the reductions in base drag and total drag can be as high as 64% an...

Unsteady compressible Reynolds-averaged Navier-Stokes equations are solved for computing the flowfield over a bulbous heat shield of a satellite launch vehicle at free stream Mach number of 0.95 and 1.20 and at zero angle of incidence. A time-dependent computation is carried out employing a multistage Runge-Kutta time-stepping in conjunction with a finite volume discretization. Closure of these equations is achieved using the Baldwin-Lomax turbulence model. Comparisons are made with the experimental results such as schlieren picture and surface pressure distribution. They are found in good agreement. Numerical analysis is used to determine the characteristics of the fluctuating surface pressure at transonic and supersonic speeds. Standard deviations, higher moments, histograms, and spectrum of pressure and sound pressure level of fluctuating pressure are analyzed in the separated region of the boattail of the heat shield. High frequency components of pressure amplitude and sound pressure levels are found to be dominate at supersonic Mach number as compared to transonic Mach number.

The global characteristics of tube train flows under wide ranges of blockage ratios and train speeds are comprehensively investigated by quasi-one-dimensional numerical simulation and theoretical analysis. The established quasi-one-dimensional model for the tube train flow and the numerical method are both proven to be reliable and efficient. Investigation reveals the typical patterns of both inviscid and viscous flows. The transition limits for these flow patterns are determined. The flows tend to be unstarted when the train moves at a high subsonic speed or a low supersonic speed and tend to be started otherwise. The transition limits between started and unstarted modes do not vary much between inviscid and viscous flows, which allows the use of analytical inviscid criteria in approximate estimation of the start/unstart state of the viscous flow. Between the supersonic isentropic limit and the Kantrowitz limit lies a dual-solution area. Which solution the flow follows depends on how the attempted cruise speed of the train is reached; a large acceleration tends to induce a started flow. The unstarted inviscid tube train flow appears to be self-similar as the train suddenly moves at a constant speed. This flow involves a precursor shock wave in front of the train and a secondary shock wave on or behind the train at the same time. By contrast, the unstarted viscous flow eventually remains a steady laminar structure that moves with the train after a sufficiently long process of development. In the stabilized flow, the precursor shock wave only occurs in the supersonic case, and the secondary shock wave only occurs in the subsonic case. For both inviscid and viscous flows, the unstarted flow can be divided into two submodes according to whether the secondary shock wave detaches from the train. Aerodynamic drag on the train is strongly correlated with the flow modes. The drag in unstarted flow increases with train speed, while the drag coefficient peaks at the secondary shock attach–detach limit. The wall transport effects cause larger drag but do not change its overall features.

This paper presents the results of an investigation of the flow around rectangular and chamfered cavities at high subsonic and low supersonic speeds. Pressures measured on the faces of the cavities are integrated to find the pressure drag of the cavities. The types of cavity tested range from simple sawcuts to cavities so long that the two ends can be regarded as independent and the results for these are compared with the sum of the drags of isolated forward and rear facing steps.Although the Reynolds numbers of the tests are similar to those in flight conditions the maximum depth of the cavities tested is only 6 mm so that the pressure resolution on the vertical faces of the cavities is limited. In spite of this it is estimated that the maximum error in the drag of any particular cavity is less than the skin friction drag on a smooth surface equal in area to half the plan area of the cavity.

The oblique wing—aircraft design for transonic and low supersonic speeds

A hybrid computational technique which splits the flowfield into inviscid and viscous regions is used to investigate the complete flowfield about axisymmetric parabolic blunt bodies in a supersonic stream. The solutions are carried out on the CDC CYBER-203 computer which, with its extensive memory, allows for the use of a large number of finite-difference mesh points, allowing resolution of important flowfield features. A range of freestream Mach number of 2-5 and a range of Re number based on nose radius of 500-125,000 was run for a sonic corner body. Contour plots of density, pressure, and Mach number, velocity vector plots, and surface distributions of pressure, heat transfer, and shear stress are presented. Also, correlations of the downstream extent of the base recirculation region with Re number based on nose radius are given.

A computational capability has been developed for predicting the flowfield about projectiles, including the recirculatory base flow at transonic speeds. In addition, the developed code allows mass injection at the projectile base and hence is used to show the effects of base bleed on base drag. Computations have been made for a secant-ogive-cylinder projectile for a series of Mach numbers in the transonic flow regime. Computed results show the qualitative and quantitative nature of base flow with and without base bleed. Base drag is computed and compared with the experimental data and semiempirical predictions. The reduction in base drag with base bleed is clearly predicted for various mass injection rates. Results are also presented that show the variation of total aerodynamic drag both with and without mass injection for Mach numbers of 0.9 < M< 1.2. The results obtained indicate that, with further development, this computational technique may provide useful design guidance for projectiles. MAJOR area of concern in shell design is the accurate prediction of the total aerodynamic drag. Both the range and terminal velocity of a projectile (two critical factors in shell design) are directly related to the total aerodynamic drag. The total drag for projectiles can be divided into three components: 1) pressure drag (excluding the base region), 2) viscous (skin friction) drag, and 3) base drag. At transonic speeds, base drag constitutes a major portion of the total drag. For a typical shell at M = 0.90, the relative magnitudes of the aerodynamic drag components are: 20% pressure drag, 30% viscous drag, and 50% base drag. The critical aerodynamic behavior of projectiles, indicated by rapid changes in the aerodynamic coefficients, occurs in the transonic speed regime and can be attributed in part to the complex shock structure existing on projectiles at transonic speeds. Therefore, in order to predict the total drag for projectiles, computation of the full flowfield (including the base flow) must be made. There are few reliable semiempirical procedures that can be used to predict shell drag; however, these procedures cannot predict the effects of mass injection. The objective of this research effort was to develop a numerical capability, using the Navier-Stokes computational technique, to compute the flowfield in the base region of projectiles at transonic speeds and thus to be able to compute the total aerodynamic drag with and without mass injection. The pressure and viscous components of drag generally cannot be reduced significantly without adversely affecting the stability of the shell. Therefore, recent attempts to reduce the total drag have been directed toward reducing the base drag. A number of studies have been made to examine the total drag reduction due to the addition of a boattail.1 Although this is very effective in reducing the total drag, it has a negative impact on the aerodynamic stability, especially at transonic