Memorandum presenting an investigation in the 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.82 and 2.01 to determine the longitudinal and lateral stability characteristics of a fighter-type airplane model with a low-aspect-ratio unswept wing and a high horizontal tail. The results indicated a region of reduced longitudinal stability at low lifts at a Mach number of 2.01 that was caused by fuselage flow fields or vertical-tail effects on the horizontal tail.

Attenuation of tangent-pressure oscillation in liquid-oxygen-heptane rocket engine combustion chamber using longitudinal fin. Results regarding the engine performance with injector A, effect of fins with injector A, and destructiveness of screaming are provided.

An investigation of the use of fins to prevent high-frequency combustion-pressure oscillations (screaming) in the combustion chamber of a 1000-pound-thrust rocket engine with a chamber pressure of 300 pounds per square inch and using liquid oxygen and n-heptane as propellants. Tangential combustion-pressure oscillations were found to be eliminated with longitudinal fins located in the combustion zone. Results regarding engine performance, effect of fins, and destructiveness of screaming are provided.

Memorandum presenting an annotated bibliography of NACA reports containing information of interest to designers and manufacturers of controls for turbojet engines. The eleven major topics include starting, acceleration, flame-out, temperature control, stability and dynamics of small disturbances, combustion dynamics, engine inlet diffuser control, steady-state engine performance, fuel systems, sensors, and general control system theory.

Report presenting an acceleration control that operates on fuel flow and uses directly measured acceleration as the control signal as applied to an axial-flow turbojet engine. A study to determine the feasibility and dynamic characteristics of the control was undertaken. Results regarding the open-loop characteristics and closed-loop characteristics are provided.

Memorandum presenting an investigation of a full-scale afterburner with high burner-inlet velocity conducted to determine burner performance with several variations in burner design. Variables receiving particular attention were flameholder design and burner length. Results regarding diffuser performance, performance of the reference two-V-gutter flameholder, designs to improve combustion for high-velocity conditions, and operational characteristics are provided.

Report presenting an investigation of a full-scale afterburner with high burner-inlet velocity to determine burner performance with several variations in burner design. A total of 12 flameholder configurations were investigated at a specified burner-inlet velocity and a range of burner-inlet pressures. Results indicated that a basic annular two-V-gutter flameholder can operate at 90 to 95 percent combustion efficiency for fairly optimum burner length and pressure.

Report presenting the stability characteristics of an axisymmetric conical inlet with supercritical spillage. Either injection or suction through the slot increased the stable subcritical range over that of the unmodified cone as much as 22 percent of the free-stream mass flow. Results regarding the effect of spike tip translation on inlet performance, effect of air injection and suction, secondary-flow requirements, comparison of various stabilizing methods, and effect of stabilizing devices on flow distortion are provided.

Free flight test of tailless missile configuration with 45-deg sweptback wing of aspect ratio 4 - stability at transonic speeds. The longitudinal flexible-wing results indicated a gradual transonic trim change a lift-curve slope comparable with that for the same configuration with a wing of aspect ratio 5.5. Results regarding trim characteristics, lift and longitudinal stability, and side-force and lateral stability are provided.

Memorandum presenting a flight test of a long-range missile configuration with a 45 degree sweptback wing of aspect ratio 4 conducted at a range of Mach and Reynolds numbers. The longitudinal flexible-wing results indicated a gradual transonic trim change and a lift-curve slope comparable with that for the same configuration with a wing of aspect ratio 5.5.

Memorandum presenting an investigation of a semispan-wing-fuselage model with an 80 degree swept pointed wing with either an unbalanced constant-chord control or an overhang-balanced swept hinge-line control in the 9- by 12-inch blowdown tunnel. Control hinge moments and effectiveness characteristics were obtained over an angle-of-attack range of plus or minus 16 degrees at control deflections up to 25 degrees. Results regarding control effectiveness, control hinge moments, control drag, and evaluation of control characteristics are provided.

Memorandum presenting a low-speed investigation to determine the effect of a fuselage on the longitudinal aerodynamic characteristics of a 45 degree sweptback wing equipped with 0.35-semispan double slotted flaps. The wing had an aspect ratio of 3.7, a taper ratio of 0.41, and a streamwise thickness ratio of 0.086.

Memorandum presenting a high-flow sonic-inlet compressor rotor employing double-circular-arc blade sections that was designed and tested in Freon-12. Results regarding overall performance, radial distribution, and blade-element performance are provided.

Report presenting a free-flight investigation over a range of Mach numbers at a range of lift coefficients to determine the aerodynamic characteristics of an airplane-configuration model with a 52.5 degree delta wing and a low, swept horizontal tail. Results regarding time history, longitudinal trim, lift, drag, longitudinal static stability, side-force characteristics, and theoretical comparisons are provided.

Memorandum presenting a free-flight investigation to determine the lift, drag, static stability, and hinge-moment characteristics of a rocket-powered model of a rocket-powered model of a ballistic-missile configuration at supersonic speeds. The model consisted essentially of a body of fineness ratio 16.9 and a cruciform set of small 60 degree delta fins located approximately 1 body diameter from the base of the body. The model lift and pitching-moment coefficients were nonlinear with angle of attack.

Report presenting a free-flight investigation to determine the lift, drag, static stability, and hinge-moment characteristics of a rocket-powered model of a ballistic-missile configuration. The purpose of this particular investigation was to determine these characteristics with a long run of boundary layer ahead of small 60 degree delta stabilizing fins.

From Introduction: "As part of a complete investigation of the J71-A-2(600-D1) turbojet engine conducted in an altitude test chamber at the NACA Lewis laboratory, the steady-state altitude performance, with afterburner inoperative and ejector shroud removed, was obtained and is presented herein. The component performance of the J71-A-2(600-D1) turbojet engine is presented in reference 1. The effects of compressor interstage bleed and adjustable inlet guide vanes on compressor-stall characteristics are described in reference 2."

Report presenting an investigation in the altitude test chamber to determine the altitude performance of the J71-A-2(600-D1) turbojet engine. The engine was equipped with two-position inlet guide vanes and compressor acceleration bleeds. Results regarding the compressor performance, combustor performance, turbine performance, and effect of altitude on engine performance are provided.

Memorandum presenting an analysis of liquid methane, ethene, and propane along with normally liquid hydrocarbon fuels. The lower molecular weight hydrocarbons are superior to current JP-type fuels as heat sinks and that the more volatile fuels may be required with cooled-turbine engines.

Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.

Report presenting the performance of a side inlet with a fixed 12 degree two-dimensional compression surface for a range of Mach numbers, angles of attack, and yaw. The effects of several methods of compression-surface boundary-layer removal were investigated as well as a solid ramp.

"An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a 1/40-sca!e pod-separation model of the Convair B-58 airplane at Mach numbers of 1.41, 1.61, and 2.01. The complete return component and the complete pod with its canard controls were tested, both alone and in the presence of each other, at Mach numbers of 1.61 and 2.01. The complete pod and various combinations of its components were tested alone at Mach numbers of 1.41, 1.61, and 2.01 and also in the presence of the return component at Mach numbers of 1.61 and 2.01" (p. 1).