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Boundary-layer transition data on a cone and flat plate obtained in the Mach 3.5 Pilot Low-Disturbance Tunnel at NASA Langley are presented. Measured flat-plate transition Reynolds numbers in this tunnel are an order of magnitude higher than previous results obtained in conventional noisy supersonic wind tunnels. Transition predictions based on compressible linear stability theory and the e N method with 7V= 10 are in excellent agreement with the measured locations of transition onset for both the cone and flat plate under low-noise conditions in this tunnel. This investigation has resolved the discrepancies between the results of linear stability theory and the data from conventional supersonic wind tunnels regarding the ratio of cone-to-flat-plate transition Reynolds numbers. Nomenclatureb -flat-plate leading-edge bluntness F -dimensionless frequency, 2irfv e /ul f -frequency M = Mach number N = exponential factor in amplification ratio ef rom linear stability theory P = pressure Pr -Prandtl number R = reference Reynolds number, (UgS/v e ) l/2 R w =freestream unit Reynolds number, p^u^/R e = local unit Reynolds number, p e u e /ii e Re T = local transition Reynolds number based on flow distance to transition r = recovery factor defined by (T aw -T e )/(T 0 -T e ) rms = root mean square s = surface distance from cone tip or leading edge of flap plate T = temperature u -streamwise velocity X -axial distance from nozzle throat or from cone tip AA^AYjAZ = axial, vertical, and horizontal dimensions of quiet test core (see Fig. 1) /3 = Mach angle H = dynamic viscosity v -kinematic viscosity, p/p p = mass density a = spatial amplification rate Subscripts aw -adiabatic wall e -local values at boundary-layer edge L = laminar o = stagnation; also onset of instability p -surface pitot tube

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T O advance boundary-layer stability and transition research and to ultimately provide reliable predictions of transition for supersonic flight vehicles, a wind tunnel is required with very low stream disturbance levels comparable to free-flight conditions. Previous investigations indicated that the freestream noise in pilot quiet nozzles is primarily caused by transition in the nozzle wall boundary layers that are subjected to Gortler instability. A new concept for nozzle design was developed that provides a large increase in the length of the quiet test core by postponing the initiation and decreasing the growth rate of Gortler vortices and thus delays transition on the nozzle walls. A new advanced Mach 3.5 axisymmetric quiet nozzle was fabricated and tested to prove the new design concept. The Reynolds numbers based on the measured length of the quiet test core for this new nozzle are in excellent agreement with the theoretical predictions. ContentsThe high noise levels in conventional supersonic wind tunnels cause premature boundary-layer transition on test models. There can be many sources for this noise, but at higher Mach numbers (>2.5) the primary source is the eddy-Machwave radiation from the turbulent boundary layer on the nozzle wall. The key requirement for the design of a quiet supersonic nozzle is to maintain the nozzle wall boundary layer in a laminar state.The primary causes of nozzle wall turbulence are 1) continuation onto the nozzle wall of the turbulent boundary layer present on the wall of the upstream piping and settling chamber, and 2) destabilization of the nozzle wall laminar boundary layer by the formation and amplification of instability waves in the nozzle wall boundary-layer flow. The turbulent boundary layer in the subsonic approach to the nozzle is removed by a suction slot upstream of the throat so that a new, laminar boundary layer will develop downstream of the slot and delay transition in the supersonic portion of the nozzle. The technique of the suction slot has been applied to all of the previous pilot quiet nozzles. 1 These pilot quiet nozzles were designed with rapid expansion contours to obtain the

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