by National Aeronautics and Space Administration, Scientific and Technical Information Office, for sale by the National Technical Information Service in [Washington], Springfield, Va .
Written in English
|Statement||Charles J. Camarda, Langley Research Center|
|Series||NASA technical paper ; 1320|
|Contributions||United States. National Aeronautics and Space Administration. Scientific and Technical Information Office, Langley Research Center|
|The Physical Object|
|Pagination||36 p. :|
|Number of Pages||36|
Get this from a library! Aerothermal tests of a heat-pipe-cooled leading edge at Mach 7. [Charles J Camarda; United States. National Aeronautics and Space Administration. Scientific and Technical Information Office.; Langley Research Center.]. The AEDC-VKF materials testing wedge (Fig. 2) was used to support the test specimens. The wedge is basically a in. by in.-long flat plate mounted on a deg wedge block. The flat plate has a in. back-step occurring in. aft of the leading edge. Since the. The methodology relating these tests and the techniques available to the experimentalist need to be clearly understood. This paper provides a brief survey of the methodology, typical facilities, and the aerothermal test techniques that have been used in the U.S. to develop hypersonic by: 2. The present contribution addresses the aerothermal experimental and computational study of a trapezoidal cross-section model simulating a trailing edge cooling cavity with one rib-roughened wall and slots along two opposite walls.
pro le shapes, leading edge diameters, Mach number distributions, and the overall pressure ratio across the vane and blade. Figure 11 highlights the heat transfer distribution associated. Effeats of Mach number, altitude end surface temperature on looal heat transfer factor "h" 7 Effeot of altitude on aem - heating rate for oonstan% Rex, M ard *w/T, 8 Variation of local heat transfer faotor h with dietonoe from nose or leeding edge for hI = , altit ft end. Heat pipes designed for transferring heat from the leading edge of hypersonic-vehicle wings, which are subjected to intense aerodynamic heating, were analyzed by Cao and Faghri . These heat pipes transfer the intense aerodynamic heating at the leading edge of the wings to a condenser where the heat is rejected by radiation or convection. Full text of "NASA Technical Reports Server (NTRS) Engineering Aerothermal Analysis for X Thermal Protection System Design" See other formats AM A A /U/S AIAA Engineering Aerothermal Analysis for X Thermal Protection System Design Kathryn E. Wurster, Christopher J. Riley, and E. Vincent Zoby NASA Langley Research Center Hampton, VA 36th .
Figure 10 shows the Mach number measured on a turbine vane and blade, and highlights; (a) the strong Mach number variations near the leading edge stagnation point, (b) accelerating flow on the pressure and suction sides immediately downstream of the leading edge, (c) region of transitional boundary layer, (d) regions of accelerating turbulent. Pre- liminary water-tunnel tests of this planform indicate that the notch flap operates as postulated. Other high-lift devices used in the design include 15 percent chord leading-edge flaps on the outboard panel, 25 percent chord trailing-edge flaps, and de- flected engine nozzles. Under flow condition of Mach 7 and dynamic pressure of bar, the waverider is seen to experience maximum temperature of about K at the end of seconds. Asymmetric loading from shock wave causes bending deformation in the waverider. Maximum deflection of 4 x m is observed at the leading edge tip after 33 seconds. The inlet tip relative Mach number is mainly determined by the inlet mass flow rate and prewhirl angle. At supersonic inlet conditions, the impeller peak efficiency drops with the increase of inlet tip relative Mach number, and there is about a point drop in the peak efficiency from 95% to % corrected speed.