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Aerodynamic characteristics of a wing with unswept quarter-chord line, aspect ratio 4, taper ratio 0.6, and NACA 65A004 airfoil section: Transonic-bump method

Description: From Introduction: "This paper presents the results of the investigation of the wing alone and of the wing-fuselage configurations employing a wing with an unswept quarter-chord line, aspect ratio 4, taper ratio 0.6, and an NACA 65A004 airfoil section parallel to the air stream. The experimental results of a wing of identical plan from having an NACA 65A006 airfoil section which was tested as part of the transonic program are presented in reference 1.
Date: May 8, 1950
Creator: Myers, Boyd C., II & Wiggins, James W.

Aerodynamic study of a wing-fuselage combination employing a wing swept back 63 degrees : effectiveness of an elevon as a longitudinal control and the effects of camber and twist on the maximum lift-drag ratio at supersonic speeds

Description: Report presenting an investigation concerned with the measurement of the characteristics of longitudinal-control devices for a wing-fuselage combination with a wing with the leading edge swept back 63 degrees. Most of the investigation was devoted to testing a 30-percent-chord, 50-percent-semispan elevon, but some used upper-surface spoilers. Results regarding the characteristics of the wing with the elevon undeflected and control-surface effectiveness are provided.
Date: May 8, 1950
Creator: Olson, Robert N. & Mead, Merrill H.

Aerodynamics of slender bodies at Mach number of 3.12 and Reynolds numbers from 2 x 10(exp 6) to 15 x 10(exp 6) 2: aerodynamic load distributions of series of five bodies having conical noses and cylindrical afterbodies

Description: Report presenting an experimental investigation to determine the aerodynamic load distributions of a series of five bodies with conical or slightly blunted noses and cylindrical afterbodies in the 1- by 1-foot supersonic wind tunnel. Pressure distributions and viscous drags were measured at Mach number 3.12 for a range of Reynolds numbers and angles of attack.
Date: May 8, 1952
Creator: Jack, John R. & Gould, Lawrence I.

Altitude operational characteristics of a prototype model of the J47D (RX1-1 and RX1-3) turbojet engines with integrated electronic control

Description: Report presenting an investigation of the altitude operational characteristics of a prototype model of the J47D turbojet engines, which includes an afterburner, a variable-area exhaust nozzle, and an integrated electronic control in an altitude wind tunnel. Results regarding compressor stall, compressor unstall, combustor blow-out during acceleration, stall and blow-out protection, acceleration characteristics, deceleration characteristics, and altitude starting characteristics, afterburner operational characteristics are provided.
Date: January 8, 1952
Creator: Bloomer, Harry E.; Conrad, E. William & Sobolewski, Adam E.

An analysis of the transonic and supersonic performance of several fixed-geometry air inlets

Description: Report presenting an analysis of the maximum power thrust-minus-drag performance of several turbojet-engine air-inlet combinations over a wide range of flight conditions. The primary objective of the analysis was to use the experimentally determined drag and total-pressure-recovery characteristics to show the range of satisfactory performance of each of several fixed-geometry inlet configurations. Results regarding the performance at 35,000 feet, inlet design considerations, speed range of efficient normal-shock inlet performance, effect of altitude on performance of the fixed-geometry system, and effect of nonstandard air on the performance of a fixed-geometry system are provided.
Date: March 8, 1955
Creator: Pendley, Robert E. & Howell, Robert R.

Analysis of turbojet-engine controls for afterburning starting

Description: From Introduction: "The object of this report is to investigate the effects of after-burner lighting on the engine behavior and control-system requirements of a controlled turbojet engine. In this report, a simulation procedure for the afterburner is developed (based on ref. 3), this afterburner simulation is coordinated with previously developed nonafterburning-engine simulation procedures (ref. 4), and the compensated interacting control systems of references 1 and 2 are used. The results of an analog investigation of the the effects of an afterburner light on several controlled turbojet-engine configurations are reported and evaluated."
Date: October 8, 1956
Creator: Phillips, W. E., Jr.

Analysis of wind-tunnel tests at low speeds of a four-engine propeller-driven airplane configuration having a wing with 40 degrees of sweepback

Description: Report presenting an investigation to determine the effects of operating propellers on the longitudinal characteristics of a four-engine tractor airplane with a 40 degree swept wing with an aspect ratio of 10. Results regarding components of the lift changes due to operating propellers, components of the pitching-moment changes due to operating propellers, effect of configuration changes on the pitching-moment characteristics of the model, stick-fixed longitudinal stability of the model, reduction of adverse effects of propellers on longitudinal stability, and propulsive characteristics are provided.
Date: October 8, 1954
Creator: Edwards, George G. & Buell, Donald A.

Analytical and experimental investigation of the effects of compressor interstage air bleed on performance characteristics of a 13-stage axial-flow compressor

Description: Air was bled over the fifth-and tenth-stage rotor-blade rows through ports designed to pass 11 and 9 percent of the inlet flow, respectively, at 80 percent speed. Along the rated operating line the maximum speed at which rotating stall was encountered was lowered by either of these bleeds, and the stall patterns below these speeds were altered so that no dangerous resonant rotor-blade bending vibrations were excited. The combination of the two bleeds completely eliminated rotating stall to at least 50 percent speed. The compressor-discharge weight flow was decreased only at intermediate speeds, and the overall pressure ratio was affected only at intermediate speeds, and the overall pressure ratio was affected only by the combination bleed at intermediate speeds. Fifth-stage bleed increased compressor efficiency at low speeds, and tenth-stage bleed decreased efficiency at intermediate speeds.
Date: February 8, 1957
Creator: Lucas, James G.; Geye, Richard P. & Calvert, Howard F.

An analytical method for evaluating factors affecting application of transpiration cooling to gas turbine blades

Description: From Introduction: "A survey of some of the advantages and problems associated with transpiration cooling of gas-turbine engines is given in reference 1, and its is shown therein that high pressure gradients around the periphery of gas-turbine blades require that the blade wall permeability be varied around the blade periphery in order for uniform cooling to be obtained over the entire blade surface. This fact is verified in experimental investigations of transpiration-cooled turbine blades mounted in a static cascade (references 2 and 3) where it is shown that although transpiration cooling results in extremely effective cooling in the midchord region of the blade, there are very large variations in the chordwise temperature distribution because of improper permeability variation."
Date: September 8, 1952
Creator: Esgar, Jack B.

An analytical study of sideslip angles and vertical-tail loads in rolling pullouts as affected by some characteristics of modern high-speed airplane configurations

Description: From Introduction: "The rolling-pullout maneuver (any maneuver in which rolls occur during high g flight conditions) has been shown to be pertinent to design considerations from the standpoint of the loads produced on a vertical tail (refs. 1 to 3). In order to understand this problem better, the results of an analytical study of effects of large variations of some of the lateral-stability-derivative coefficients on the maximum angle of sideslip at first peak of its oscillation are presented in this paper. Also, expressions for estimating the effects of small variations of or errors in these coefficients on the maximum sideslip angle have been developed and are presented. "
Date: October 8, 1953
Creator: Stone, Ralph W., Jr.

Characteristics of perforated diffusers at free-stream Mach number 1.90

Description: "An investigation was conducted at Mach number 1.90 to determine pressure recovery and mass-flow characteristics of series of perforated convergent-divergent supersonic diffusers. Pressure recoveries as high as 96 percent were obtained, but at reduced mass flows through the diffuser. Theoretical considerations of effect of perforation distribution on shock stability in converging section of diffuser are presented and correlated with experimental data. A method of estimating relative importance of pressure recovery and mass flow on internal thrust coefficient basis is given and a comparison of various diffusers investigated is made" (p. 1).
Date: May 8, 1950
Creator: Hunczak, Henry R. & Kremzier, Emil J.

Chemical igniters for starting jet fuel - nitric acid rockets

Description: Report presenting experiments in a 200-pound-thrust jet fuel-nitric acid rocket with an axial-flow igniter employing the hypergolic alkyl thiosphosphite-nitric acid system. No combustion instability was observed in the igniter, but many of the main-stage runs exhibited low-frequency oscillations.
Date: October 8, 1957
Creator: Morrell, Gerald & Ladanyi, Dezso J.

Comparison between prediction and experiment for all-movable wing and body combinations at supersonic speeds: Lift, pitching moment, and hinge moment

Description: From Summary: "A simple method is presented for estimating lift, pitching-moment, and hinge-moment characteristics of all-movable wings in the presence of a body as well as the characteristics of wing-body combinations employing such wings. In general, good agreement between the method and experiment was obtained for the lift and pitching moment of the entire wing-body combination and for the lift of the wing in the presence of the body. The method is valid for moderate angles of attack, wing deflection angles, and width of gap between wing and body."
Date: August 8, 1952
Creator: Nielsen, Jack N.; Kaattari, George E. & Drake, William C.

Comparison of airfoil sections on two triangular-wing-fuselage configurations at transonic speeds from tests by the NACA wing-flow method

Description: Report presenting tests using the NACA wing-flow method at a range of Mach numbers to determine the aerodynamic characteristics of four triangular-wing-fuselage models. Measurements of the normal force, chord force, and pitching moment for various angles of attack. Two models had wings of aspect ratio 2.31, one with an NACA 65-009 airfoil section and one with a 9-percent-thick biconvex section, and two with wings of aspect ratio 4, one with an NACA 65-006 airfoil and one with a 6-percent-thick double-wedge airfoil.
Date: August 8, 1951
Creator: Hall, Albert W. & McKay, James M.

A comparison of flight-measured carrier-approach speeds with values predicted by several different criteria for 41 fighter-type airplane configurations

Description: Report presenting lift and drag characteristics in the landing-approach configuration on 41 jet-propelled fighter-type airplane arrangements using flight testing. Minimum comfortable approach speeds were evaluated by four test pilots. Suggestions for achieving optimal speeds for approach under various conditions are also provided.
Date: May 8, 1958
Creator: White, Maurice D.; Schlaff, Bernard A. & Drinkwater, Fred J., III

Coolant-Flow Calibrations of Three Simulated Porous Gas-Turbine Blades

Description: An investigation was conducted at the NACA Lewis laboratory to determine whether simulated porous gas-turbine blades fabricated by the Eaton Manufacturing Company of Cleveland, Ohio would be satisfactory with respect to coolant flow for application in gas-turbine engines. These blades simulated porous turbine blades by forcing the cooling air onto the blade surface through a large number of chordwise openings or slits between laminations of sheet metal or wire. This type of surface has a finite number of openings, whereas a porous surface has an almost infinite number of smaller openings for the coolant flow. The investigation showed that a blade made of sheet-metal laminations stacked on a support member that passed up through the coolant passage was completely unsatisfactory because of extremely poor coolant flow distribution over the blade surface. The flow distribution for two wire-wound blades was more uniform, but the pressure drop between the coolant supply pressure and the local pressure on the outside of the blades was too low by a factor ranging from 3 to 3.5 for the required coolant flow rates. The pressure drop could be increased by forcing the wires closer together during blade fabrication.
Date: March 8, 1951
Creator: Esger, Jack B. & Lea, Alfred L.

Correlation of vibratory root failures and stress distribution in J65 compressor blades

Description: Report presenting an investigation of the stress distribution in the roots of the first three stages of the J6S axial-flow compressor in order to explain root failures experienced in service. Root failures were producible in the lab when high vibratory stresses were combined with simulated centrifugal loads. Results regarding fatigue tests on the original J6S blade roots and on two redesigned blade roots as well as a comparison of the different root designs are provided.
Date: September 8, 1955
Creator: Meyer, André J., Jr. & Kaufman, Albert

Damping in Pitch of Low-Aspect-Ratio Wings at Subsonic and Supersonic Speeds

Description: Memorandum presenting the application of the concept of indicial functions to the analysis of the aerodynamic phenomena associated with the short-period pitching mode of wings in subsonic and supersonic flight. Simple physical relationships are pointed out and are used to study the effect on the rotary-damping-moment coefficient of changes in center-of-gravity position, Mach number, aspect ratio, plan form, frequency, and thickness.
Date: April 8, 1953
Creator: Tobak, Murray

Damping in roll of rocket-powered test vehicles having swept, tapered wings of low aspect ratio

Description: From Summary: "Flight tests of rocket-powered models have been conducted to determine the damping in roll of a group of swept, tapered wings designed for flight in the transonic speed region. The Mach number range of these tests was from approximately 0.7 to 1.4. The experimental damping in roll for all configurations was less than predicted by linearized theory throughout the Mach umber range of these tests. The only wing in this group that experienced an appreciable transonic lateral trim change was the one with a 7-percent-thick circular-arc airfoil section."
Date: October 8, 1951
Creator: Sanders, E. Claude, Jr. & Edmondson, James L.

Derivation of the equations of motion of a symmetrical wing-tip-coupled airplane configuration with rotational freedom at the junctures

Description: From Summary: "The method of Lagrange multipliers is used to take account of the dynamic effects of the constraints at the wing tips when two identical airplanes are coupled to the wing tips of a "mother" airplane. The resulting equations of motion of this symmetrical configuration are derived for one, two, or three degrees of rotational freedom at each joint."
Date: October 8, 1951
Creator: Schy, Albert A.