Last edited by Tugal
Tuesday, July 28, 2020 | History

3 edition of Aerothermodynamic calculations on X-34 at Mach 6 wind tunnel conditions found in the catalog.

Aerothermodynamic calculations on X-34 at Mach 6 wind tunnel conditions

William A. Wood

Aerothermodynamic calculations on X-34 at Mach 6 wind tunnel conditions

by William A. Wood

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  • 33 Currently reading

Published by National Aeronautics and Space Administration, Langley Research Center, National Technical Information Service, distributor in Hampton, Va, [Springfield, Va .
Written in English

    Subjects:
  • X-34 reusable launch vehicle.,
  • Wind tunnel tests.,
  • Laminar boundary layer.,
  • Hypersonic heat transfer.,
  • Reynolds number.,
  • Turbulence effects.,
  • Aerothermodynamics.,
  • Turbulent boundary layer.,
  • Baldwin-Lomax turbulence model.,
  • Aerodynamic heating.

  • Edition Notes

    StatementWilliam A. Wood.
    SeriesNASA/TM -- 1999-208998., NASA technical memorandum -- 208998.
    ContributionsLangley Research Center.
    The Physical Object
    FormatMicroform
    Pagination1 v.
    ID Numbers
    Open LibraryOL15542540M

    Aerothermodynamics. Flow of gases in which heat exchanges produce a significant effect on the flow. Traditionally, aerodynamics treats the flow of gases, usually air, in which the thermodynamic state is not far different from standard atmospheric conditions at sea level. And so on the shuttle we didn't really do much aerothermodynamic testing beyond mach eight. That gave us the right normal shock Reynolds number region, it was closer to, on the Shuttle, the actual flight environment than if we had gone to a m a m a m a mach 20 facility.

      Propulsion Wind Tunnel 16T. The facility has two foot by foot long test section, closed-circuit wind tunnels. One is transonic (16T), one is supersonic (16S). 16T can be operated from Mach numbers to 16S is currently inactive, however, when it was operational it could operate from Mach numbers to The Langley 8-Foot High Temperature Tunnel (8-Ft HTT) is a combustion-heated hypersonic blowdown-to-atmosphere wind tunnel that provides simulation of flight enthalpy (relationship of an object, its environment, and energy) for Mach numbers of 4, 5, and 7 and Reynolds numbers from x 10 6 to x 10 6 per foot, depending on the Mach number.

    An experiment analyzing performance of wind tunnel and co-relating theoretical analysis is presented. Various losses at different sections in tunnel have been calculated and total pressure drop across the wind tunnel has been estimated using them. Accordingly, total power required or total power invested to run a opencircuit wind. studied in the NASA Langley Mach quiet-tunnel under low-noise conditions comparable to °ight [17]. Measurements were made on an inch °at plate, 11 inches from the leading edge. The bluntness was small and the wall was adiabatic. Transition onset was measured at Re. x = 18£ 6: Measurements of the end of transition at Re. x ’


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Aerothermodynamic calculations on X-34 at Mach 6 wind tunnel conditions by William A. Wood Download PDF EPUB FB2

Around Mach 6, with turbulent flow. NASA Langley has participated in a task agreement with Orbital Sciences to assist in the prediction of the X aerothermodynamic environment and the analysis of the aerothermal design of the vehicle. To this end, a series of wind tunnel experiments and numerical simulations have been performed.

Berry et. Get this from a library. Aerothermodynamic calculations on X at Mach 6 wind tunnel conditions. [William A Wood; Langley Research Center.].

Aerothermodynamic Calculations on X at Mach 6 Wind Tunnel Conditions. By William A. Wood. Abstract. The effects of Reynolds number and turbulence on surface heat-transfer rates are numerically investigated for a scale X vehicle at wind tunnel conditions.

Laminar heating rates, non-dimensionalized by Fay-Riddell stagnation heating Author: William A. Wood. Inviscid Flow Computations of the Orbital Sciences X Over a Mach Number Range of to by National Aeronautics and Space Administration NASA.

Aerothermodynamic Calculations on X at Mach 6 Wind Tunnel Conditions. by National Aeronautics and Space Administration NASA. Book Depository Books With Free Delivery Worldwide. operating conditions considered is estimated to be less than 10% of the rotor thrust.

Wind Tunnel Test Results. The test results reported here consist of four runs (described in table 1), at nominal longitudinal separation distances of x/D =, and (two runs at x/D = ). The wind tunnel Cited by: Research Center: the Inch Mach 6 Air Tun-nel, the Inch Mach 10 Air Tunnel, the Inch Mach 6 CF4 Tunnel, and the Inch Mach 20 Helium Tunnel.

Table 1 gives a summary of the nominal reservoir and corresponding free-stream flow conditions in each facility along with. The Inch Mach 6 Tunnel was used to acquire data denoted as T and the Inch Mach 10 Tunnel was used to acquire data denoted T, T, and T They are described detail in Micol Martian space probe at wind tunnel conditions, corresponding to Mach numbers M =5 and 14 for various angles of attack is carried out on basis of Navier-Stokes equations, which are solved using in-house software package HSFlow with possibility of distributed computing on multiprocessor computer.

Peculiarities in behavior of aerodynamic and. Most of the NASA's recently proposed X-vehicles have been tested in the NASA LaRC Inch Mach 6 Air Tunnel with a majority of the aerothermodynamic studies emphasizing hypersonic transition and.

destroyed during reentry. Aerothermodynamic calculations must guarantee that the burn-up of the vehicle in the atmosphere takes place the vehicle are: – ground-based facilities such as classical wind tunnels, shock tunnels, plasma facilities and their instrumentation – numerical analysis codes, ranging from simple and fast engineering tools.

Inch Mach 6 Tunnel The Langley Inch Mach 6 Tunnel, which became operational inis a blowdown wind tunnel that uses dry air as the test gas. Air from two high pressure bottlefields is transferred to a psia reservoir and is heated within this reservoir to a maximimi temperature of °R by an electrical resistance heater.

The wind tunnel facilities provide continuous high enthalpy flows which aim to replicate the aerothermodynamic environment of the post shock stagnation streamline. The goal of such experiments is to simulate the stagnation point heat load by replicating the boundary layer thermochemical state [8].

The usefulness of this quality measure is assessed by comparing heat transfer predictions from grid convergence studies for grids of varying quality in the range of [] on an 8 half angle sphere-cone, at laminar, perfect gas, Mach 10 wind tunnel conditions.

The maximum model scale for the Mach 6 tests in Tunnel B was Smaller facilities, specifically the Langley m (inch) Mach 6 Wind Tunnel and the m (inch) Mach 10 Wind Tunnel, were initially excluded from stage-separation tests, because their relatively small size could not accommodate a free flyer and the full-length.

whereby data from hypersonic wind-tunnel tests are integrated with predictions from CFD X, X, X, and X as well as to the Space Shuttle Columbia accident investigation and return to flight While most of the data presented in this overview paper were measured in the Mach 6.

When wind tunnel testing, you must closely match the Mach number between the experiment and flight conditions. It is completely incorrect to measure a drag coefficient at some low speed (say mph) and apply that drag coefficient at twice the speed of sound (approximately mph, Mach = ).

In Fig. 1 1 the MINIVER engineering predictions for the windward centerline of a scale model of the X config- uration X at a wind-tunnel condition of Mach 6, an AOA of 15 deg, and unit Reynolds number of million are compared to thermographic phosphor data 32 and analytical predictions.

', - 25 In Fig. 1 1 (a) the MINIVER. The speed of sound in air at room temperature is approximately m/s and for this case Mach number is less than one (Mwind tunnels, the maximum speed achieved by the wind tunnel is equal to or greater than the speed of sound in air hence Mach number is greater than 1(M>1).

wind tunnel results are often available for the low speed regime only. Therefore, in many to deliver aerothermodynamic data up to flight Mach numbers between 7 and 8 at altitudes between 90 and 20 km.

under well controlled flow conditions – not necessarily exactly the same as fully realistic flight conditions.

However, useful. used in defining wind-tunnel test conditions, in interpreting the measured data, and finally for the flight extrapolation.

Figure 4 shows some interesting flow patterns on the windward and lee sides of the X vehicle at Mach and 40 deg incidence. It confirms the predicted (Lore) increased radiation equilibrium temperatures at the. The maximum model scale for the Mach 6 tests in Tunnel B was Smaller facilities, specifically the Langley m (inch) Mach 6 Wind Tunnel and the m (inch) Mach 10 Wind Tun­ nel, were initially excluded from stage-separation tests, because their relatively small size could not accommodate a free flyer and the full-length.Hypersonic aerodynamics is the most active frontier in the field of fluid mechanics.

Its development is closely allied to the development of modern hypersonic flight vehicles. The future development of hypersonic aerodynamics can be expected to include: establishing predictive models within abound physical framework; developing high-performance computing and managing capabilities to handle.vehicles, the X small reusable technology demonstrator program.

Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on and scale models of proposed X flight vehicles at Mach 6 and 10 in air.

The primary parametrics that were investigated include angles-of-attack from 0.