Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows

Detailed aerodynamic heating measurements were made on a 70-deg sphere-cone configuration model and in the wake of the model on the sting. Tests were conducted in hypersonic flows in a high-enthalpy impulse facility, in which air and carbon dioxide were employed as test gases, and in a conventional perfect-gas air wind tunnel. Heating data were also obtained on three similar parametric forebody configurations. Normalized forebody Stanton number distributions were independent of Reynolds number and test gas, with the exception of smaller forebody corner heating peaks in carbon dioxide. Peak wake Stanton numbers were 5% of the forebody stagnation point values in the high-enthalpy tests and varied with Reynolds number from 7 to 15% of the stagnation point values in the perfect-gas tests. The impulse facility wake flow establishment process was studied in detail, and a criterion for determining when the wake flow becomes established was developed. Wake flow establishment was found to require on the order of 45-75 flow-path lengths as based on the model size and freestream velocity. NomenclatureC = ^Too/^ooT-* C H = Stanton number, q/[pooUoo(ho -h w )] c p = specific heat, J/kg-K h = enthalpy, J/kg k = thermal conductivity, W/m-K L = surface distance along sting from model base, m Q -heat energy, J/m 2 q = heat transfer rate, W/m 2 . Re -Reynolds number, pUx/iJi R h = forebody base radius, m 5 = surface distance measured from the stagnation point, m T = temperature, K 7* = reference temperature, (7o, 2 /6)[T+ (3T w /T Qt2 )], K t = time, s U -velocity, m/s y K f = reference length for flow establishment a = thermal diffusivity, k/pc p , m 2 /s ft = thermal product, j(kpc p ) 9 W-s 1/2 /m 2 -K A = heat transfer correction factor, 1/K A = uncertainty bound for a parameter fji = viscosity, kg/m-s jL6* = viscosity evaluated at 7*, kg/m-s p = density, kg/m 3 a = heat transfer residual for flow establishment T = nondimensional flow establishment time, U^ At/y nf Subscripts w = wall conditions 0= total or stagnation conditions 0, 2 = post-normal-shock stagnation conditions 2 = post-normal-shock static conditions oo = freestream conditions

Section: Comparison Of Mach 10 and Hypulse Resultsmentioning

confidence: 78%

High-enthalpy and perfect-gas heating measurements on a blunt cone

Hollis

Perkins

1996

“…This was clearly evident on the rst manned hypersonic test vehicle, the X-15, where on a ight in which a ramjet test article was attached to the ventral n, structural damage to the engine pylon occurred as a result of shock impingement and the test article was lost in ight. Designers of subsequent hypersonic ight vehicles,such as the Shuttle Orbiter, have, at some point in the Reno, NV, Jan [15][16][17][18]. 1996; received Feb. 29, 1996; revision received March 10, 1997; accepted for publication March 15, 1997.…”

mentioning

confidence: 99%

Fin Leading-Edge Sweep Effect on Shock-Shock Interaction at Mach 6

Berry

Nowak

1997

The effects of n leading-edge sweep on peak heating rates due to shock-shock interaction have been experimentally examined in the NASA Langley Research Center 20-Inch Mach 6 Tunnel. The shock interaction was produced by the intersection of a planar incident shock (16.8-deg shock angle relative to the freestream, generated by a 9-deg wedge) with the bow shock formed around a 0.5-in.-diam cylindrical leading-edge n. Heating distributions along the leading-edge stagnation line have been obtained using densely spaced thin-lm resistive-type sensors. Schlieren images were obtained to illustrate the very complex shock-shock interactions. The n leading-edgesweep angle was varied from 15 deg swept back (relative to the normal to the freestream) to 45 deg swept forward for a freestream unit Reynolds number of 2 £ 10 6 /ft. Two models were utilized during the study, one with 0.025-in. spacing between gauge centers, and the other with 0.015-in. spacing. Gauge spatial resolution on the order of 0.015 in. appeared to capture the narrow heating spike accurately. Peak heating due to shock interaction was maximized when the n was swept forward 15 and 25 deg, both promoting augmentationsover seven times the baseline value. The schlieren images for these cases revealed Type IV and Type III interactions, respectively. NomenclatureCh = nondimensional Stanton number, q=½ 1 u 1 .h t2 ¡ h 1 ) C h ref = reference Stanton number, measured for each run from undisturbed end of model h t2 = stagnation enthalpy behind normal shock, Btu/lbm h 1 = freestream enthalpy, Btu/lbm L = length of leading edge, 4 in. M 1 = freestream Mach number P t1= stagnation pressure, psi q = heat transfer rate, Btu/ft 2 ¢ s Re 1 /ft = freestream unit Reynolds number, ft ¡1 T t 1 = stagnation temperature, ± R T 1 = freestream temperature, ± R t = time, s u 1 = freestream velocity, ft/s x = distance measured from incident shock location, in. = incident shock angle relative to the freestream, deg ± = shock generator angle relative to the freestream, deģ = n sweep angle relative to the normal to the freestream (where positive is de ned as swept back and negative as swept forward), deg ½ 1 = freestream density, slug/ft 3

“…The freestream-based parameters are simpler to determine and do tend to correlate the data from a single facility. Additionally, the edge-based definitions contain assumptions that apply only to a blunt-body stagnation region and would not be applicable to a flat-plate, small-angle sphere-cone, or complex geometry; see for example the correlation for small-angle biconics in [26], which includes freestream Mach number and cone-angle corrections. Thus, to avoid ambiguity, windtunnel data are typically reported in terms of freestream conditions such as shown in Figs.…”

Section: Application Of Laminar Heating Distribution Correlationmentioning

confidence: 99%

Blunt-Body Entry Vehicle Aerotherodynamics: Transition and Turbulent Heating

Hollis

2012

Recent, current, and planned NASA missions that employ blunt-body entry vehicles pose aerothermodyamic problems that challenge state-of-the-art experimental and computational methods. The issues of boundary-layer transition and turbulent heating on the heat shield have become important in the designs of both the Mars Science Laboratory and Crew Exploration Vehicle. While considerable experience in these general areas exists, that experience is mainly derived from simple geometries; e.g., sharp-cones and flat-plates, or from lifting bodies such as the Space Shuttle Orbiter. For blunt-body vehicles, application of existing data, correlations, and comparisons is questionable because an all, or mostly, subsonic flowfield is produced behind the bow shock, as compared with the supersonic (or even hypersonic) flow of other configurations. Because of the need for design and validation data for projects such as Mars Science Laboratory and Crew Exploration Vehicle, many new experimental studies have been conducted in the last decade to obtain detailed boundary-layer transition and turbulent heating data on this class of vehicle. In this paper, details of several of the test programs are reviewed. The laminar and turbulent data from these various test are shown to correlate in terms of edge-based Stanton and Reynolds number functions. Correlations are developed from the data for transition onset and turbulent heating augmentation as functions of momentum thickness Reynolds number. These correlations can be employed as engineering-level design and analysis tools.adiabatic wall enthalpy H 0 = total freestream enthalpy H w = wall static enthalpy H 300 K = wall static enthalpy at 300 K M 1 = freestream Mach number M e = boundary-layer edge Mach number P 1 = freestream pressure Pr = Prandtl number _ q w = heat transfer rate R N = hemispherical nose radius Re 1 = freestream unit Reynolds number 1 U 1 = 1 Re e = boundary-layer edge unit Reynolds number e U e = e Re = boundary-layer momentum thickness Reynolds number e U e = e St = Stanton number based on freestream conditions _ q w = 1 U 1 H 0 H w St e = Stanton number based on boundary-layer edge conditions _ q w = e U e H aw H w T 1 = freestream temperature U 1 = freestream velocity U e = boundary-layer edge velocity x=R = normalized distance along model centerline for Mars Science Laboratory local coordinate system z=R = normalized distance along model centerline for Crew Exploration Vehicle local coordinate system = angle of attack L = factor in laminar heating correlation 1 1 = e e 1=2 e = 1 1=4 T = factor in turbulent heating correlation 1 1 = e e 1=3 e = 1 1=4 = boundary-layer momentum thickness 1 = freestream viscosity e = boundary-layer edge viscosity = turbulent heating augmentation factor St measured =St predicted;laminar I. BackgroundT HIS work deals with the correlation of blunt-body boundarylayer transition and turbulent heating data obtained from recent NASA entry vehicle development programs. Blunt-body configurations are the most common geometries employed for ent...