Future Prospects for Hypersonic Flight Test For University Students

The transport to space of satellites and experiments with conventional expandable rockets is a very expensive and time demanding business and therefore almost prohibitive for Universities and High-School or Colleges laboratories. In this lecture an unconventional, low-cost reusable concept-system based on "electromagnetic railgun " is presented and discussed. The aims is to provide in the midterm with a system capable of launch in LEO small satellites of maximum 5 kg weight not by using conventional chemical power but accelerating them up to hypersonic speed with help of electromagnetic forces. In the short term such system could offer the possibility to carry out low-cost hypersonic- and atmospheric-research experiments. The subject of the lecture is to demonstrate the feasibility of launching small meteorological science experiment-payloads by means of hypersonic projectiles, which are accelerated with an electromagnetic railgun up to an end-velocity of about 2100m/s. Associated with this new non- conventional propulsion technique is the development of hypersonic projectiles which shall withstand a harsh thermo-mechanical environment resulting from accelerations of about 10,000g. Compared to the solid and/or liquid rockets available today, a railgun launcher concept-system offers several advantages such as high efficiency with high repetitions rates andlow recurring costs than today's market prices

Designing Flight Experiments for Hypersonic Flow Physics

A major problem of concen for the success of physical-modelling resides in the availability of experimental data for model validation, particularly in the hot hypersonic regime. In the past, validation data have been achieved as secondary product of expensive space-transportation programs. Since in the last ten years there has been almost no successfull program due to lack of investment, no new experimental data are available. According, a new trend is emerging for low cost technology validation in flight based on the use of old military rockets. Here it is described the possibilities of sounding rockets to provide a platform for flight experiments in hypersonic conditions as cost-efficient supplement of tests in ground based facilities. The lecture regards the overall philosophy of projects flown on sounding rockets, but mainly the current efforts of DLR on the SHarp Edge Flight EXperiment SHEFEX. The approaches adopted on each of the involved disciplines, i.e. mission system and launcher; aerodynamics, aerothermodynamics and in-flight measurement techniques; structure and thermal protection systems are discussed. The paper is aimed to show that even flying hypersonic is not a routine technology, the strategy proposed is not only a highly efficient way to acquire important knowledge in the physics of hypersonic flows but also an excellent test bed for proving new technological concepts

Compressible Inviscid Vortex Flow of a Sharp Edge Delta Wing

An analysis is presented on the compressible inviscid vortex flow over three delta wings with sharp leading edges with leading edge sweep angles *&#6160, 70, and 76 deg using numerical solutions of the Euler equations. Presentations of results are given for Mach numbers ranging from M*&#610.1 to 0.8 and for angles of attack up to the onset of vortex breakdown. The paper focuses the attention on the effects of the vortex flow on the flowfield of a delta wing. The occurence of shock waves of two types, crossflow and terminating (or rear) shocks in the vortical flowfields, are investigated; the possibility of shock-induced vortex breakdown and the effects of compressibility on the rapid performance degradation of delta wings after vortex bursting are studied in detail. It is shown that the onset of the vortex bursting at the wing trailing edge is only weakly influenced by the freestream Mach number

German Participation in the Aerothermodynamic Design of the Experimental Vehicle X-38

The experimental vehicle X-38 is a technology demonstrator for the future Crew Rescue Vehicle of the International Space Station. The program, leads by the US space agency NASA, is the first cooperation for the design of a reusable space vehicle between NASA, the European Space Agency ESA and the German space agency DLR. The paper briefly describes the main objective of the program and its technological goals with emphasis on the German participation. Particularly, the work-methodology adopted by the three agencies to build up the aerodynamic database and to establish the aerothermal environments of the experimental vehicle is presented and discussed. It is shown that the selected functional-scheme successfully work and can be adopted for future transatlantic projects

Sensitivity Study for the Simulation of the ARD Capsule under Transonic Flow Conditions

Sensitivity studies for the transonic simulation of the ARD capsule without and with the reaction yaw control system activated are presented and discussed. The motivation of this investigation lies in the large discrepancies in pressure level found in preliminary studies when comparing numerical results with measured values. Flow simulations including the wind tunnel support, several algebraic models, variable angle of attack and variable amount of numerical dissipation are performed. The computed results are found independent of the type of turbulence models here tested. Including the wind tunnel support in the numerical simulation the pressure level close to the capsule backcover reduces, increasing mainly the axial force. The amount of numerical dissipation used in the computed solution has the largest effect in the pressure level. By reducing the numerical dissipation the computed pressure levels approach to that of the

Flap Efficiency versus Flap Heating of a Winged Reentry Vehicle- A Challenge for the CFD-

In the present study Navier-Stokes solutions are analyzed for a winged reentry vehicle at hypersonic free stream conditions corresponding to both, cold wind tunnel tests and a flight trajectory point. The flow was computed as laminar, assuming calorically perfect gas for the wind tunnel case. The high-temperature effects on thermodynamic and transport properties are modelled by assuming air in thermochemical equilibrium. Computations are presented for two types of control surfaces geometry: cambered and plain, for clean configurations and for configurations with deflected control surfaces. The accuracy of the computed solutions is addressed by grid refinement studies and by comparing numerical results with available experimental data. Also the importance of the viscous effects is shown by comparing the Navier-Stokes solutions with Euler solutions

The Compressible Inviscid Vortex Flow of a Sharp Delta Wing

An analysis is presented on the compressible inviscid vortex flow over three delta wings with sharp leading edges with leading edge sweep angles 60, 70 and 76 degrees using numerical solutions of the Euler equations. Presentations of results are given for Mach numbers ranging from M &#61 0.1 to 0.8 and for angles of attack up to the onset of vortex breakdown. The paper focusses the attention on the effects of the vortex flow on the flow field of a delta wing. The occurence of shock waves of two types: cross-flow and terminating (or rear) shocks induced vortex breakdown and the effects of compressibility on the rapid performance degradation of delta wings after vortex bursting are studied in detail. Moreover, the changes on flow kinematic within the vortex core due to the close presence of the wing are also studied

A comparison of some zero equation and half equation turbulence models for hypersonic flow

The objective of the present research is to validate several zero equation -algebraic- and half equation turbulence models on prediction of the hypersonic flow properties. The computations are performed by solving the Reynolds-averaged Navier-Stokes equations with the turbulent viscous terms closed by 6 alternative eddy-viscosity models due to Baldwin and Lomax; Launder and Priddin; Granville; Johnson and King; Johnson and Abid and Chima et al. the numerical algorithm used is the DLR computer code CEVCATS. It is based on a finite volume method coupled with a multi-stage Runge-Kutta scheme. Solutions are obtained for the body flap of the reentry vehicle HERMES at conditions corresponding to cold wind tunnel -M*&#6110, *&#6130*, *&#6110* - and compared with available experimental data. No attempts is done to model transition.Instead off, turbulent flow is enforced at the point where a laminar flow solution begins to diverge from the experimental data

Modelling of Hypersonic Flow Phenomena

This study focuses on the numerical simulation of supersonic flows around a missile with lattice wings. In order to reduce the computational cost for such a configuration, the actuator disc concept is chosen to model the effects of grid fins. This approach is coupled with a Navier-Stokes solver in order to predict the forces and moments on a complete vehicle. The method consists in replacing the lattice controls by artificial boundary conditions where the forces involved by the grid fins are applied. These forces are interpolated from an experimental database. Numerical simulations are performed for laminar and turbulent flows for several Mach numbers and angles of attack. The results are compared to experimental data in order to validate the method. The comparisons reveal some discrepancies which are mainly due to the turbulence effects and to the database from which the forces resulting from the lattice wings are interpolated. The method allows the prediction of the forces applied on the vehicle and therefore an estimate of its aerodynamic performances. The computations validate the approach and show its potential as a tool for vehicle design