Verification and Analysis of Formulation 4 of Langley for the Study of Noise From High Speed Surfaces

There are several approaches to the prediction of the noise from sources on high speed surfaces. Two of these are the Kirchhoff and the Ffowcs williams-Hawkings methods. It can be shown that both of these methods depend on the solution of the wave equation with mathematically similar inhomogeneous source terms. Two subsonic solutions known as Formulation 1 and 1A of Langley are simple and efficient for noise prediction. The supersonic solution known as Formulation 3 is very complicated and difficult to code. Because of the complexity of the result, the computation time is longer than the subsonic formulas. Furthermore, it is difficult to assess the accuracy of noise prediction. We have been searching for a new and simpler supersonic formulation without these shortcomings. In the last AIAA Aeroacoustics Conference in Toulouse, Farassat, Dunn and Brentner presented a paper in which such a result was presented and called Formulation 4 of Langley. In this paper we will present two analytic..

Scott C. Asbury and Jeffrey A. Yetter NASA Langley Research Center Hampton, Virginia

An experimental investigation was conducted in the Jet-Exit Test Facility at NASA Langley Research Center to study the static aerodynamic performance of a wing-mounted thrust reverser concept applicable to subsonic transport aircraft. This innovative engine powered thrust reverser system is designed to utilize wing-mounted flow deflectors to produce aircraft deceleration forces. Testing was conducted using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0, a supercritical left-hand wing section attached via a pylon, and wing-mounted flow deflectors attached to the wing section. Geometric variations of key design parameters investigated for the wing-mounted thrust reverser concept included flow deflector angle and chord length, deflector edge fences, and the yaw mount angle of the deflector system (normal to the engine centerline or parallel to the wing trailing edge). All tests were conducted with no external flow and high pressure air was used to s..

Lawrence D. Huebner, Kenneth E. Rock, Randall T. Voland, and Allan R. Wieting NASA Langley Research Center Hampton, VA

The NASA Langley 8-Foot High Temperature Tunnel has recently been modified to produce a unique testing capability for hypersonic airbreathing propulsion systems. Prior to these modifications, the facility was used primarily for aerothermal loads and structural verification testing at true flight total enthalpy conditions for Mach numbers between 6 and 7. One of the recent modifications was an oxygen replenishment system which allows operating airbreathing propulsion systems to be tested at true flight total enthalpies. Following the modifications to the facility, calibration runs were performed at total enthalpies corresponding to flight Mach numbers of 6.3 and 6.8 to establish the flow characteristics of the facility with its new capabilities. The results of this calibration, as well as modifications to tunnel combustor hardware prior to calibration to improve tunnel flow quality, are described in this paper. Introduction The NASA Langley 8-Foot High Temperature Tunnel (8 HTT) was de..

X-38 Experimental Aeroheating at Mach 10

This report provides an update of the hypersonic aerothermodynamic wind tunnel test program conducted at the NASA Langley Research Center in support of the X-38 program. Global surface heat transfer distributions were measured on 0.0177 and 0.0236 scale models of the proposed X-38 configuration at Mach 10 in air. The parametrics that were investigated primarily include freestream unit Reynolds numbers of 0.6 to 2.2 million per foot and body flap deflections of 15, 20, and 25 deg. for an angle-of-attack of 40-deg. The model-scale variance was tested to obtain laminar, transitional, and turbulent heating levels on the deflected body flaps. In addition, a limited investigation of forced boundary layer transition through the use of discrete roughness elements was performed. Comparisons of the present experimental results to computational predictions and previous experimental data were conducted. Laminar, transitional, and turbulent heating levels were observed on the deflected body flap, which compared favorably to the computational results and to the predicted heating based on the flight aerothermodynamic database. NOMENCLATURE h heat transfer coefficient (lbm/ft 2 -sec), = q/(H aw -H w ) where H aw = H t2 H enthalpy (BTU/lbm) k trip height (in) L reference length taken from nose to end of body M free stream Mach number P pressure (psia) q heat transfer rate (BTU/ft 2 -sec) Re unit Reynolds number (1/ft) T temperature (R) t time (sec) angle of attack (deg) BF Body flap deflection (deg) Subscripts free-stream conditions t1 reservoir conditions t2 stagnation conditions behind normal shock w wal