Recent advances in the aerothermodynamics of spiked hypersonic vehicles

Drag Reduction Using Aerodisks for Hypersonic Hemispherical Bodies

Forebody shock control devices for drag and aero-heating reduction: A comprehensive survey with a practical perspective

A spiked blunt body flying at hypersonic speed could result in significant thermal interactions between the external flowfield and the internal structure, which have not received enough attention in previous investigations. In the present study, time-dependent, loosely coupled fluid–thermal interaction simulations are conducted on a blunt body equipped with a spike with variable length at a M=5, Re=2×106 flight condition. The results show that, in the range of time considered in the study, a blunt body with a longer spike yields a better dynamic thermal performance, that is, a lower peak temperature on the blunt forebody surface and a slower temperature rise both on the forebody surface and inside the structure can be obtained. In addition, various levels of drag coefficient drop can be observed during the coupling process. A maximum decrease of 5.7% is achieved among the investigated models. The study reveals the necessity of the coupled fluid–thermal analysis to accurately predict the aerothermal enviro...

Fluid–Thermal Interaction Investigation of Spiked Blunt Bodies at Hypersonic Flight Condition

Among various design requirements, reducing both drag and aerodynamic heating is the most crucial in the design of hypersonic vehicles. Pointed slender bodies generate lower drag compared with blunt bodies at hypersonic conditions. Hence, they provide a good choice as far as drag reduction is concerned. On the other hand, blunt bodies were found to yield lower heating levels compared with their pointed counterparts [1]. In fact, hypersonic vehicles such as missiles, interplanetary space missions, space planes, and launch missiles usually have blunt shapes [1],[2]. Obviously, designing a vehicle that simultaneously satisfies both minimum drag and aerodynamic heating is not straightforward and there is a challenging trade-off between these two vital requirements. However, it is believed that these two requirements can be met by altering the flowfield pattern around the blunt body so as to eliminate the strong detached bow shock wave. A variety of techniques have been implemented in this regard [1] but the use of spikes proved to be the simplest and the most effective technique in reducing both drag and aerodynamic heating [3].

Surrogate-Based Multi-Objective Aerothermodynamic Design Optimization of Hypersonic Spiked Bodies

Many design concepts of high-speed tactical guided missiles use aerospikes to reduce the drag and heating on the missiles body. The accurate numerical prediction of hypersonic flows over aerospike-protected missile shapes using Computational Fluid Dynamics (CFD) can be a valuable tool in the evaluation of such co ncepts. This study presents an effective and robust numerical approach to the problem of a hypersonic flow over a missile body shielded with aerospike/aerodisc. The problem considers three-dimensional flow around aerospikeprotected missile domes at a free stream Mach number ranging from 4.5 to 6.06 and angles of attack varying from 0 to 16 deg. Steady-state nu merical simulations are carried out using the general purpose CFD code FLUENT, version 6.3. Pressure-based coupled solver formulation is chosen to obtain an accurate fast-co nverging solution. SST ω k turbulence model is employed to simulate turbulence effects. T he near-wall mesh is fine enough to resolve the viscosity-affected near-wall region all the way to the laminar sublayer, and FLUENT’s enhanced wall treatment is applied. Numerical results expressed in terms of pressure coefficient are compared with experimental data, and favorable agreement is found. Effects of aerospike on missile dome surface aeroheating are also presented.

Application of Pressure-Based Coupled Solver to the Problem of Hypersonic Missiles with Aerospikes

Numerical Blast Wave Identification and Tracking Using Solution-Based Mesh Adaptation Approach

Specialized computer models have been developed by others, to predict the multiphase flow through SRM nozzles, as well as the rocket plu me. In this study, two approaches are discussed for calculating the effects of a particul ate phase on the nozzle and plume flow in the continuum regime. The focus of the work is the comparison of computational fluid dynamics (CFD) models in which the particulate phase is calculated in either a Langrangian or Eulerian reference frame, and the discussion of the multiphase nozzle and plume flow dynamics observed. Steady-state, compressible, non-reacting, turbulent, multiphase flow is assumed. Phase change and radiative heat transfer are not considered. The properties, size distribution, and mass loading for the particulate phase are assumed and kept constant. A conical nozzle profile is chosen with assumed inlet flow conditions (including a 30% particulate mass loading), and the multiphase nozzl e flow is compared with a solution obtained using a specialized computer model. The plume is then expanded into an initially ambient environment, and the multiphase plume flow resulting from the CFD solutions on two different computational grids is discussed. Co mplex mean flow effects were modeled and shown to have great importance in compressible multiphase nozzle and plume flow simulations. The methodologies presented are general and can be applied to 3D and parallel computations.

Multiphase CFD Analysis of Solid Rocket Motor Nozzle and Plume Flow

Many design concepts of unmanned aerial vehicles (UAV), especially mini- and microUAVs, use oscillating and flapping wings for propulsion and control. The accurate numerical prediction of transient turbulent flows over airfoils with complex motion via CFD can be a valuable tool in the evaluation of such concepts. This study presents and compares two moving mesh methods to analyze unsteady turbulent flows over an oscillatory pitching airfoil. The problem considers two-dimensional flow around NACA64A010 airfoil at a free stream Mach number of 0.796. This flow is transonic with weak shocks developing near the mid-chord. Unsteady numerical simulation in the time domain is carried out using the general purpose CFD code FLUENT, version 6.2. The realizable e k turbulence model is employed to simulate turbulence effects. The near-wall mesh is fine enough to resolve the viscosity-affected near-wall region all the way to the laminar sublayer, and Fluent’s enhanced wall treatment is applied. The oscillatory pitching motion of the airfoil around its quarter-chord is prescribed using two separate approaches: (i) the sliding mesh model where the moving zone encompassing the airfoil oscillates in discrete steps relative to the surrounding stationary zone with both zones communicating via a non-conformal sliding interface, and (ii) the dynamic mesh model where the mesh surrounding the airfoil is deformed based on the new positions of the airfoil using the spring-based smoothing method. Both approaches reproduce numerical solutions in close agreement with each other. The lift and moment coefficients are compared with experimental data and favorable agreement is found. Other turbulence models including Spalart-Allmaras, SST ω k and F v 2 are also considered.

Konstantin A. Kurbatskii

Papers

Application of Pressure-Based Coupled Solver to the Problem of Hypersonic Missiles with Aerospikes

Numerical Blast Wave Identification and Tracking Using Solution-Based Mesh Adaptation Approach

Multiphase CFD Analysis of Solid Rocket Motor Nozzle and Plume Flow

Application of Two Moving Mesh Methods to the Problem of Unsteady Transonic Flow Over Oscillatory Pitching Airfoil