Static internal performance of a nonaxisymmetric vaned thrust reverser with flow splay capability

An investigation was conducted in the Static Test Facility of the Langley 16-Foot Transonic Tunnel on a dual-port, nonaxisymmetric, block-and-turn type thrust reverser model with vane cascades in the reverser ports which turned the flow in the splay direction and aided in turning the flow in the reverse direction. Splaying reverser flow is a method of delaying to lower landing ground roll speeds the reingestion of hot exhaust flow into the inlets. Exhaust flow splay can also help prevent the impingement of hot exhaust gases on the empennage surfaces when the reverser is integrated into an actual airframe. The vane cascades consisted of two sets of perpendicular vanes with a variable number of turning and splay vanes. A skewed vane box was also tested which had only one set of vanes angled to provide both turning and splay. Vane cascades were designed to provide different amounts of flow splay in the top and bottom ports. Inner doors, trim tabs, and an orifice plate all provided means of varying the port area for reverser flow modulation. The outer door position was varied as a means of influencing the flow reverse angle. Nozzle pressure ratio was varied from 1.75 to approximately 6.00

Experimental Sonic Boom Measurements on a Mach 1.6 Cruise Low-Boom Configuration

A wind tunnel test has been conducted by Gulfstream Aerospace Corporation (GAC) to measure the sonic boom pressure signature of a low boom Mach 1.6 cruise business jet in the Langley Unitary Plan Wind Tunnel at Mach numbers 1.60 and 1.80. Through a cooperative agreement between GAC and the National Aeronautics and Space Administration (NASA), GAC provided NASA access to some of the experimental data and NASA is publishing these data for the sonic boom research community. On-track and off-track near field sonic boom pressure signatures were acquired at three separation distances (0.5, 1.2, and 1.7 reference body lengths) and three angles of attack (-0.26deg, 0.26deg, and 0.68deg). The model was blade mounted to minimize the sting effects on the sonic boom signatures. Although no extensive data analysis is provided, selected data are plotted to illustrate salient features of the data. All of the experimental sonic boom pressure data are tabulated. Schlieren images of the configuration are also included

Effect of Afterbody Geometry on Aerodynamic Characteristics of Isolated Nonaxisymmetric Afterbodies at Transonic Mach Numbers

A parametric study was conducted in the Langley 16-Foot Transonic Tunnel on an isolated nonaxisymmetic fuselage model that simulates a twin-engine fighter. The effects of aft-end closure distribution (top/bottom) nozzle-flap boattail angle versus nozzle-sidewall boattail angle) and afterbody and nozzle corner treatment (sharp or radius) were investigated. Four different closure distributions with three different corner radii were tested. Tests were conducted over a range of Mach numbers from 0.40 to 1.25 and over a range of angles of attack from -3 to 9 degrees. Solid plume simulators were used to simulate the jet exhaust. For a given closure distribution in the range of Mach numbers tested, the sharp-corner nozzles generally had the highest drag, and the 2-in. corner-radius nozzles generally had the lowest drag. The effect of closure distribution on afterbody drag was highly dependent on configuration and flight condition

Report Documentation Page

(Maximum 200 words) A parametric study has been conducted in the Langley 16-Foot Transonic Tunnel on an isolated nonaxisymmetric fuselage model that simulates a twin-engine fighter. The effects of aft-end closure distribution (top/bottom nozzle-flap boattail angle versus nozzle-sidewall boattail angle) and afterbody and nozzle corner treatment (sharp or radius) were investigated. Four different closure distributions with three different corner radii were tested. Tests were conducted over a range of Mach numbers from 0.40 to 1.25 and over a range of angles of attack from \Gamma3 ffi to 9 ffi . Solid plume simulators were used to simulate the jet exhaust. For a given closure distribution in the range of Mach numbers tested, the sharp-corner nozzles generally had the highest drag and the 2-in. corner-radius nozzles generally had the lowest drag. The effect of closure distribution on afterbody drag was highly dependent on configuration and flight condition. 14. SUBJECT TERMS 15. NUMBE..

Historical Overview, Facility Description, Calibration, Flow Characteristics, and Test Capabilities

The Langley 16-Foot Transonic Tunnel is a closed-circuit single-return atmospheric wind tunnel that has a slotted octagonal test section with continuous air exchange. The wind tunnel speed can be varied continuously over a Mach number range from 0.1 to 1.3. Test-section plenum suction is used for speeds above a Mach number of 1.05. Over a period of some 40 years, the wind tunnel has undergone many modifications. During the modifications completed in 1990, a new model support system that increased blockage, new fan blades, a catcher screen for the first set of turning vanes, and process controllers for tunnel speed, model attitude, and jet flow for powered models were installed. This report presents a complete description of the Langley 16-Foot Transonic Tunnel and auxiliary equipment, the calibration procedures, and the results of the 1977 and the 1990 wind tunnel calibration with test-section air removal. Comparisons with previous calibrations showed that the modifications made to the..

Research Opportunities in Advanced Aerospace Concepts

This report is a review of a team effort that focuses on advanced aerospace concepts of the 21st Century. The paper emphasis advanced technologies, rather than cataloging every unusual aircraft that has ever been attempted. To dispel the myth that "aerodynamics is a mature science" an extensive list of "What we cannot do, or do not know" was enumerated. A zeit geist, a feeling for the spirit of the times, was developed, based on existing research goals. Technological drivers and the constraints that might influence these technological developments in a future society were also examined. The present status of aeronautics, space exploration, and non-aerospace applications, both military and commercial, including enabling technologies are discussed. A discussion of non-technological issues affecting advanced concepts research is presented. The benefit of using the study of advanced vehicles as a tool to uncover new directions for technology development is often necessary. An appendix is provided containing examples of advanced vehicle configurations currently of interest