TITLE AUTHORS CONTRACT- REPORT M. D. Nelson, L. L. Linscheid, B. A. Dinwiddie, III, and O. J. Hall, (NASA-CR-1853: D3-8535) Avail: NTIS CSCL 018 A study and development program was accomplished to attenuate turbine noise generated in the JT3D turbofan engine. Analytical studies were used to design an acoustic liner for the tailpipe. Engine ground tests defined the tailpipe environmental factors and laboratory tests were used to support the analytical studies. Furnace-brazed, stainless steel, perforated sheet acoustic liners were designed, fabricated, installed, and ground tested in the tailpipe of a JT3D engine. Test results showed the turbine tones were suppressed below the level of the jet exhaust for most far field polar angles. PATENT CITATION Author SOURCE PUBLICATION DATE AVAILABILITY COSATI PLACE OF PRIMARY ACCESSION AND LOCATION OF ABSTRACT Arthur P. Kubicz, inventor (to NASA) Issued 15 Jun. 1971 8 p (NASA-Case-GSC-10667-1; US-Patent-3,585,514; PATENT APPLICATION CITATION N72-10432# National Aeronautics and Space Administration. DIATOMIC INFRARED GASDYNAMIC LASER Patent Robert L. McKenzie, inventor (to NASA) Filed 26 Apr. 1971 U.S. PATENT NO. U.S. PATENT APPLICATION (NASA-Case-ARC-10370-1; US-Patent-Appl-SN-137391) Avail: NTIS CSCL 20E PRIMARY ANALYTIC CITATION AND ABSTRACT FULL CITATION AND N72-10856*# National Aeronautics and Space Administration. THE GEMINI PROGRAM: PHYSICAL SCIENCES Edward O. Zeitler. comp. and Thomas G. Rogers, comp. (ITT Results and findings of physical sciences experiments SUBSIDIARY ANALYTIC CITATION N72-10862*# Geological Survey, Flagstaff, Ariz. Elliott C. Morris In NASA. Manned Spacecraft Center The Avail: NTIS HC $6.00/MF $0.95 CSCL 14E SERIAL NO. EXCEPTION SUBSIDIARY ANALYTIC ACCESSIONS PAGES IN REPORT PRICE OF PRIMARY REPORT VOLUME 10 NUMBER 4 FEBRUARY 23, 1972 Scientific and Technical Aerospace Reports A Semimonthly Publication of the National Aeronautics and Space Administration ADMINISTRATION = Avail: NTIS = Analytical and experimental phases of the subject investigation are described. The analytical program for the single jet determines the terminal shock location, the jet boundary, the interface profile, the bow shock profile, the shear layer growth and the dead air region pressure. The experimental program described was conducted over the range from free stream Mach 0.4 to 2.0 at angles-of-attack up to 18 deg and at thrusting coefficients up to C sub T T/q sub infinity A sub m 30. Variables investigated included aeroshell angle, number of nozzles, engine thrust, size of nozzles, nozzle throttling and gas composition. The influence of these variables on the aeroshell stability, drag, and loads was determined by integrating pressure measurements on the aeroshell. The total system forces consist of components due to pure thrust and components due to pressure on the aeroshell arising from the jet-free stream interaction. Shadowgraphs provided flow field geometries which proved to be within 10% of those predicted analytically. Author DIGITAL COMPUTER CALCULATIONS OF THE EFFECTS OF VARIATION OF MASS AND INERTIA PARAMETERS ON THE SPIN OF A 60 DEG DELTA-WING FIGHTER AIRCRAFT R. Wilson Jul. 1971 46 prefs (WRE-TN-453-(WR/D); COSATI-0103; COSATI-2004; RD73) Avail: NTIS Spinning motions of a 60 deg delta-wing fighter aircraft were obtained by numerically integrating the six degrees of freedom equations of motion on a digital computer. The effects of variation of mass and inertia parameters were studied by comparing perturbed parameter spins with a datum spin while the aerodynamic properties were kept constant. Large and small parameter variations were considered separately, with the small variations being compatible with practical uncertainties in estimates of aircraft mass and inertias, and the large variations being compatible with the carriage of external stores. The effects on computed spins of small parameter variations were not significant. Of the large parameter variations, increasing the rolling moment of inertia was found to be the most practical means of improving spin recovery characteristics. Author N72-12975*# National Aeronautics and Space Administration. Lewis Research Center, Cleveland, Ohio. COMPUTER PROGRAM FOR DESIGN OF TWODIMENSIONAL SUPERSONIC TURBINE ROTOR BLADES WITH BOUNDARY-LAYER CORRECTION Louis J. Goldman and Vincent J. Scullin Washington Dec. 1971 58 p refs (NASA-TM-X-2434; E-6494) Avail: NTIS CSCL 20D A FORTRAN 4 computer program for the design of two-dimensional supersonic rotor blade sections corrected for boundary-layer displacement thickness is presented. The ideal rotor is designed by the method of characteristics to produce vortex flow within the blade passage. The boundary-layer parameters are calculated by Cohen and Reshotoko's method for laminar flow and Sasman and Cresci's method for turbulent flow. The program input consists essentially of the blade surface Mach number distribution and total flow conditions. The primary output is the corrected blade profile and the boundary-layer parameters. Author The side force and normal force characteristics of a spinning cylindrical body with a tangent ogive nose were determined for angles of incidence from 0 to 30 deg, spin ratios (peripheral velocity/cross flow velocity) 0 to 1, body lengths 4.42 to 8.25 kalibers and Reynolds numbers (based on diameter) of 200,000 and 250,000. The side force varies nonlinearly with spin ratio and incidence, but is always positive. Vortex locations were obtained in the wake behind the 7.07 caliber body at angles of incidence 15 to 30 deg, and spin ratios 0 to 1 in the Reynolds number range 6400 to 250,000. Evidence of tertiary vortices was not obtained except at a Reynolds number of 50,000. At a spin ratio of 0.5 the vortices assumed markedly different asymmetric patterns at Reynolds numbers of 50,000 and 200,000. Cross flow separation angles were measured along the 7.07 caliber body at angles of incidence 15 and 30 deg, spin ratios 0 to 1 and Reynolds numbers 50,000 and 200,000. The cross flow separation position is found to be weakly dependent on axial position. The cross flow separation angle moves with the spin except on the high pressure side of the model at a spin ratio bf 0.5 and a Reynolds number of 200,000. Author N72-12978# Advisory Group for Aerospace Research and Development, Paris (France). AERODYNAMIC TESTING AT HIGH REYNOLDS NUMBERS AND TRANSONIC SPEEDS D. Kuechemann (Roy. Aircraft Estab., Farnborough, Engl.) Nov. 1971 9 p refs Presented at NATO Defence Res. Group Seminar on Gen. Probl. Relating to Aerodyn. Testing Facilities, St. Louis, France, 4-7 May 1971 (AGARD-R-588-71) Avail: NTIS As a contribution to a NATO seminar on aerodynamic testing facilities held at the ISL from 4 to 7 May 1971, a brief report is given on the outcome of a specialists' meeting organized by the fluid dynamics panel of AGARD and held at Gottingen from 26 to 28 April 1971. Various AGARD activities, leading up to this meeting, are also briefly described. The best technical advice available within AGARD leads to the conclusion that one or several large new wind tunnels would contribute immensely to the effectiveness of a large number of aerospace systems now planned or contemplated within the NATO nations. Author N72-12979*# National Aeronautics and Space Administration. Lewis Research Center, Cleveland, Ohio. TOLERANCE OF MACH 2.50 AXISYMMETRIC MIXEDCOMPRESSION INLETS TO UPSTREAM FLOW VARIA TIONS David A. Choby Washington Jan. 1972 52 prefs An investigation of the tolerances of two Mach 2.50 axisymmetric mixed-compression inlets to upstream flow variations was conducted. Tolerances of each inlet to angle of attack as a function of decreasing free-stream Mach number were obtained. A local region of overcompression was formed on the leeward side of the inlet at maximum angle of attack before unstart. This region of overcompression corresponded to local subsonic flow conditions ahead of the geometric throat. A uniform Mach number gradient of 0.10 at the cowl lip plane did not affect the inlet's pressure recovery, mass flow ratio, or diffuser exit total-pressure distortion. Author the aerodynamic characteristics of a straight wing space shuttle orbiter at subsonic speeds. A 0.02-scale model was tested at Mach numbers from 0.3 to 0.9 and Reynolds numbers from about 600,000 to 3 million, based on body width. The body alone and the body with its wing and horizontal tail attached were tested at angles of attack from 35 to 75 degrees. The effects of rounding the body corners at the junctures connecting the bottom and sides were investigated for corner radii from 0 to 8.5 percent of the body width. At low subsonic Mach numbers (free stream Mach number approximately equal 0.3) the aerodynamic characteristics are affected significantly by changes in Reynolds number and body corner radius. With increase in Mach number to free stream Mach number 0.9 the effect of Reynolds number seems to vanish, but a significant effect of body corner radius remains. Author N72-12982*# Scientific Translation Service, Santa Barbara, P. Perrier and W. Vitte Washington NASA Dec. 1971 33 p (NASA-TT-F-14074) Avail: NTIS CSCL 01A The three-dimensional problem of vortex interaction and its effect on lift, drag and other aerodynamic properties of various wing and aircraft configurations is treated by a new analytical method. The aerodynamic surfaces are replaced by equivalent surface elements. The data are displayed on an IBM terminal. N72-12983*# Author Scientific Translation Service, Santa Barbara, Calif. Y. Semezis and J. Gombert Washington NASA Dec. 1971 69 P refs Transl. into ENGLISH of "L' Aile aux Portances Elevees. Remarques sur la Prevision des Caracteristiques" Chatillon, Soc. Natl. Ind., Aerospatiale, 1971 85 p Presented at 7th Appl. Aerodyn. Colloq., Modane and Ecully, France, 4-6 Nov. 1971 (Contract NASW-2035) (NASA-TT-F-14073) Avail: NTIS CSCL 01A Problems encountered in predicting the characteristics of high lift wings, with or without additional high lift devices, are examined. Existing means of controlling lift or distributing the load on a wing by detachment are considered. Particular emphasis is placed on the processes of prediction of maximum lift coefficient and its increase by means of high lift devices. These are compared for the case of a moderately long and moderately swept-back wing equipped with conventional high lift devices. It is considered that improvement in prediction is related to a better understanding of detechment phenomena. Results of current studies are briefly reviewed, and various procedures to increase lift by blowing are described. Author |