Ivan E. Beckwith

Transition on swept leading edges at Mach 3.5

Data and correlations for transition from laminar to turbulent flow on 45and 60-deg swept cylinders are presented. The data were obtained at Mach 3.5 in the Pilot Low-Disturbance Wind Tunnel at NASA Langley. Freestream noise levels were varied during the test program from extremely low values that were essentially in the instrument noise range to much higher values approaching those in conventional wind tunnels. The results show that end plates or large trips near the upstream end of the cylinders cause turbulent flow along the entire attachment line of the models over the freestream test Reynolds number range (based on cylinder diameter) of approximately 1.0 X10 </?00jZj < 1.6 x 10. When all end disturbance sources are removed, transition occurs on the attachment lines at R^D — 7-8 x 10 independent of freestream noise levels and in agreement with previous correlations. With the addition of small roughness elements on the cylinder attachment lines, transition occurs at lower values of the Reynolds number, depending on both the roughness height and the wind-tunnel noise level. Limited data obtained off the attachment line indicate that the effects on transition of end disturbances, roughness, and wind-tunnel noise are generally similar to those on the attachment lines.

Design and fabrication requirements for low-noise supersonic/hypersonic wind tunnels

A schematic diagram of the new proposed Supersonic Low Disturbance Tunnel (SLDT) is shown. Large width two dimensional rapid expansion nozzles guarantee wide quiet test cores that are well suited for testing models at large angle of attack and for swept wings. Hence, this type of nozzle will be operated first in the new proposed large scale SLDT. Test results indicate that the surface finish of pilot nozzles is critical. The local roughness Reynolds number criteria R sub k is approx. = 10 will be used to specify allowable roughness on new pilot nozzles and the new proposed tunnel. Experimental data and calculations for M = 3.0, 3.5, and 5.0 nozzles give N-factors from 6 to 10 for transition caused by Goertler vortices. The use of N is approx. = 9.0 for the Goertler instability predicts quiet test cores in the new M = 3.5 and M = 6.0 axisymmetric long pilot nozzles that are 3 to 4 times longer than observed in the test nozzles to date. The new nozzles utilize a region of radial flow which moves the inflection point far downstream and delays the onset and amplification of the Goertler vortices.

Effects of Nozzle Design Parameters on the Extent of Quiet Test Flow at Mach 3.5

To 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 much lower stream disturbance levels than in conventional supersonic tunnels. Recent results from a Mach 3.5 pilot quiet tunnel at the NASA Langley Research Center have shown that very low stream noise levels can be achieved only when the nozzle wall boundary layers are laminar. Transition Reynolds numbers measured on a slender cone were then in the same range as flight data. The linear amplification of both Tollmien-Schlichting (TS) and Gortler type instabilities was calculated for the wall boundary layers in this Mach 3.5 pilot nozzle. By using an eN method, it was determined that amplified Gortler vortices were involved in the transition process. The resulting transition criteria were then applied to several nozzles with different expansion rates and wall curvatures. By manipulation of these parameters, the calculations indicate that it may be possible to increase significantly the size of the quiet test regions compared with those observed in the pilot nozzle.

Advanced Mach 3.5 Axisymmetric Quiet Nozzle

To 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. Experimental and theoretical research to develop a low-disturbance supersonic wind tunnel has achieved a breakthrough. A new concept for nozzle design is presented which promises a large increase in the length of the quiet test core. The Advanced Mach 3.5 Axisymmetric Quiet Nozzle is the first prototype built to prove the new design concept. Experimental results from this new nozzle on the extent of laminar wall boundary layers are compared with data from other nozzles and with theoretical predictions based on linear stability theory. 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. The effect of surface finish on the nozzle performance is also discussed.

Gortler instability and supersonic quiet nozzle design

To 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

Development of square nozzles for supersonic low-disturbance wind tunnels

Two Mach 2.4 nozzles with square test sections have been designed and analyzed, as part of an effort to develop low-disturbance facilities for laminar-flow control for high-speed civil transport. The mean flows have been simulated using a finite volume, central-differencing scheme to solve the thin-layer NavierStokes equations. Substantial crossflow exists in the boundary layers of both nozzles. The crossflow changes direction about halfway between the throat and exit, because of a change in the sign of the crossflow pressure gradient. S-shaped crossflow profiles are present in this region, just as on a swept wing. Both the standard crossflow Reynolds number and the new Reed and Haynes transition estimator predict transition to occur in the throat as well as near the exit. An analysis of the crossflow pressure gradients in the transonic throat region indicates that streamwise curvature has only a weak effect on the crossflow pressure gradient. Further research on crossflow-induced transition will have to be carried out before these nozzles will be suitable for low Mach number quiet-tunnel designs.

On the Design of a New MACH 3.5 Quiet Nozzle

To 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. A new concept for nozzle design is presented which promises a large increase in the length of the quiet test core. The Advanced Mach 3.5 Axisymmetric Quiet Nozzle is the first prototype built 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.

Transition Research in Low-Disturbance High-Speed Wind Tunnels

The technical requirements and test data from the Mach 3.5 Pilot Low-Disturbance Tunnel are presented. This unique facility provides a test region with essentially zero-acoustic noise and simulates, for the first time, the low-disturbance conditions of atmospheric flight. Applications to the test results of linear stability theory with the eN method indicate that transition locations for both simple and complex flows are well predicted by using N ≅ 9 to 11.

High Speed Boundary Layer Transition on a Blunt Nose Flare with Roughness

A blunt nose flare model was designed and constructed and is to be tested in the Langley Mach 6 quiet tunnel. The experiment is designed to investigate the effects of roughness and pressure gradient on hypersonic boundary layer transition. The design of the model is discussed, as are several new experimental techniques which are to be incorporated. Inviscid boundary layer, and stability computations are reported at both Mach 6 and Mach 3.5. Some preliminary experimental results at Mach 3.5 are also given.

Boundary-Layer Instability Mechanisms on a Swept-Leading Edge at Mach 3.5

Correlations have been made in NASA Langleys Mach 3.5 Pilot Quiet Tunnel for the transitions occurring from laminar to turbulent flow, in the cases of 45-deg and 60-deg swept cylinders. While freestream noise variations had no effect on boundary layer transition, the addition of boundary layer trips to the leading edges led to transition at lower Re numbers, depending on both trip height and wind tunnel noise level. Also presented are the results of compressible linear stability calculations for the boundary layer of an infinite swept cylinder; Tollmien-Schlichting waves are found to be amplified in the attachment line boundary layer.