A one-half-scale heat-pipe-cooled leading edge model was designed and fabricated to verify feasibility and performance of a full-scale Shuttle-type orbiter design. Model performance was investigated experimentally by radiant heating and aerothermal tests and analytically by using a simple technique which predicts heat pipe startup from the frozen state and also predicts transient and steady-state thermal behavior. Analytical results agree well with experimental results for startup and steady-state heat pipe operation. The results indicate that variations in angle of attack and roll orientation had a negligible effect on heat pipe performance. The heat pipes effectively isothermalized the leading edge, and reduced peak temperatures to levels compatible with the use temperatures of superalloys. Results of these tests demonstrated the durability of the heat-pipe-cooled leading edge in withstanding Earth-entry thermal and mechanical loads and indicate that the use of heat pipes in leading edge structures is a reliable concept for fully reusable hypersonic cruise and space transportation systems. a D So h I I T -M m P R Rn T r* 7 X Nomenclature speed of sound inside diameter of heat pipe gravitational constant latent heat of vaporization chordwise surface distance measured from nose total chordwise surface distance measured from nose to end of leading edge Mach number molecular mass of working fluid pressure heat flux sonic heat transport rate universal gas constant Reynolds number temperature temperature above which continuum flow conditions are assumed to exist in the vapor space time temperature difference angle of attack angle of roll ratio of specific heats mean free path viscosity density P = Subscripts c =combustor / = liquid max = maximum s = stagnation t = total v = vapor oo = free stream

The results of an experimental program to evaluate heat transfer and pressure distributions on corrugation roughened flat plates in thick turbulent boundary layers are presented. The experimental program consisted of tests in the tunnel wall boundary layers of the Langley Unitary Plan Wind Tunnel (UPWT) and Continuous Flow Hypersonic Tunnel (CFHT) at freestream Mach numbers of 2.5, 3.5, 4.5, and 10.3. Tests in the UPWT were conducted at a freestream Reynolds number/m of 10.8 x 10 6 and in the CFHT, at Reynolds numbers/m of 1.3 to 5.8 x 10 6 . The test configurations consisted of 50.8 cm x 50.8 cm panels with corrugated beads of two different peak amplitudes, 0.61 cm and 0.29 cm. The angle of the corrugated beads relative to the flow direction was varied between 0° (aligned) and 90° (normal). The measured peak and average heat transfer are analyzed and correlated in terms of the bulk boundary layer, internal boundary layer, and geometric parameters. The data are also compared with similar data for thinner boundary layers, and with previously published correlation techniques. Nomenclature a -speed of sound Cf = skin friction coefficient C p = specific heat h = heat transfer coefficient, q/(T aw -T w ) k e = equivalent sand grain roughness height L = wave length in direction of flow M = Mach number N St = Stanton number, h/p^U^C p P = static pressure q = heat transfer rate Re = Reynolds number S = surface distance from top of wave T = temperature U = velocity a = local wave surface angle 7 = ratio of specific heats, 7 = 1.4 e = maximum wave amplitude from wave midline d = boundary-layer thickness 6* = boundary-layer displacement thickness 6 5 = laminar sublayer thickness 6 = momentum thickness 6 = temperature ratioX7Y € -T w )/(T Too -T w ) [L = molecular viscosity 4> = angle of corrugations relative to freestream flow direction p = density Subscripts av = average aw = adiabatic wall i FP = flat plate ke -based on equivalent sand grain roughness height max = maximum Presented as Paper 75-190 at the AIAA 13th Aerospace Sciences Meeting,

The uncertainties in heating predictions derived from ground test data correlations have been used to define the corresponding uncertainties in TPS weight for the Space Shuttle. A completely reusable Shuttle system consisting of an aluminum heat sink booster and orbiter with reusable surface insulation for thermal protection was evaluated. The largest contribution to the uncertainty in TPS weight for the orbiter occurred on lower surface areas as a result of heating and boundary-layer transition uncertainties. Extension of this work to the current Shuttle system concept showed reduced weight uncertainty for the external tank compared to the reusable booster. Nomenclature H(T), h, #(0.9TO), H = heat-transfer coefficient L m M P 4 Re ReJL Re e T V W X a 6 a t A Subscripts B C.L. e L tr * oo s Ref = model reference length (body axial length) = meter = Mach number = pressure = heating rate = Reynolds number = freestream unit Reynolds number = momentum thickness Reynolds number = temperature = velocity = TPS weight = axial distance = angle of attack = local flow deflection angle = peripheral angle = population standard deviation = time = increment in parameter prefixed = booster alone = booster mated with orbiter = lower surface centerline value -edge condition = local conditions = transition location = local peripheral value = freestream = stagnation point = reference value of h or q for a unit (R = I') sphere scaled to model size = wall condition = impingement region = undisturbed Presented as Paper 73-737 at

FIGURE 3-4 BASIC WAVE CONSTRUCTION AND NOMENCLATURE 3-5 ORIGINAL •1>A~~OF pOOR QU direction. The first complete cycle and the last complete wave cycle were also instrumented. However, only the data in the center of the panel are of interest to the present study since they have previously been shown to be representative of all the waves in very thick turbulent boundary layers (neglecting edge effects).Details of the geometry and instrumentation for panels -4 and -5, and -6 are discussed in Ref. 8 and 10. respectively.

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