Memorandum presenting an investigation to determine the energy transfer by thermal radiation from the flame of a single tubular turbojet-engine combustor to the combustor liner. The effects of variations in combustor inlet-air pressure, fuel-air ratio, and air mass flow on the transfer of radiant energy were studied.

Memorandum presenting a linear theoretical analysis of the performance of three proportional navigation guidance systems installed in a given supersonic, variable-incidence, boost-glide, antiaircraft missile at Mach numbers of 2.7 and 1.3. It is shown that, with the antenna stabilized in space, the effect of component lags on the response is small, so that the speed of response is small, so that the speed of response can be made to approach closely that of the airframe alone.

Report presenting the horizontal-tail load measurements obtained during testing of the Bell X-5 airplane at the NACA High-Speed Flight Research Station. Measurements were taken at a range of Mach numbers and from sweep angles of 20 degrees to 59 degrees.

An experimental investigation of thermal radiation from the flame of a single tubular turbojet-engine combustor to the combustor liner is presented. The effects of combustor inlet-air pressure, air mass flow, and fuel-air ratio on the radiant intensity and the temperature and emissivity of the flame are reported. The total radiation of the "luminous" flames (containing incandescent soot particles) was much greater (4 to 21 times) than the "nonluminous" molecular radiation.

Results of tests to determine the aerodynamic forces and moments on bodies of revolution at angles of attack from 0 degrees to 25 degrees are presented and compared with theory. Cones and ogives of fineness ratios 3 to 7 and two blunt-nosed body shapes with fineness ratios 3 and 5 were tested at Mach numbers from 2.75 to 5.00. Reynolds numbers were from 0.5 million to 6.4 million, depending on Mach number and body fineness ratio.

Memorandum presenting an investigation conducted in a modified turbojet engine to determine the cooling effectiveness of an air-cooled turbine-rotor blade employing a corrugated metal sheet brazed to the inner surface of the blade shell to increase the internal heat-transfer area. The corrugated fin blade had a lower integrated average wall temperature at the section investigated than the more promising shell-supported blade types previously investigated. Results regarding the temperature distribution investigation and pressure loss investigation are provided.

Memorandum presenting a study of the effect of free methyl radicals on the slow oxidation of both ethane and propane by means of a photochemical decomposition reaction in a static system. The introduction of methyl radicals into a mixture of propane and oxygen or ethane and oxygen substantially lowered the initiation temperature of the combustion of the hydrocarbon.

"A preliminary investigation of the flow fluctuations of surge and blade row stall was conducted with three single-stage axial-flow compressors with hub-tip ratios of 0.9,0.8, and 0.5 and with a multistage axial-flow compressor. Flow fluctuations of large amplitude associated with stall were detected in all compressors investigated. The fluctuations were caused by low flow regions affecting 25 to 40 percent of the annulus area and propagating in the direction of compressor rotation, but at a lower speed. Mild audible surge was obtained with the single-stage compressors with hub-tip ratios of 0.9 and 0.8" (p. 1).

Performance characteristics of a two-dimensional isentropic diffuser have been experimentally determined at a Mach number of 3.85. At zero angle of attack, a maximum total-pressure recovery of 0.41 was obtained with a supercritical mass-flow ratio of 0.95. As a consequence of the twin-duct arrangement of the diffuser, a large discontinuity in pressure recovery and mass flow with a characteristic hysteresis was encountered between critical and subcritical operation. An asymmetric shock pattern with large-scale separation and flow reversal in one of the passages occurred at reduced mass flows. Pressure and force data presented for an angle-of-attack range from zero to 4 degrees.

Report presenting an investigation of a series of rectangular wing and body combinations at Mach numbers 1.62, 1.93, and 2.41 to determine the effects of aspect ratio, incidence angle, and forebody length on the interference lift, drag, and pitching moment. A limited investigation to determine the effect of Reynolds number on the wings in the presence of the body was also carried out. Information regarding wing lift, basic quantities for interference evaluation, interference qualities, contributions of the various components, and the concept of wing-lift carry-over is provided.

Memorandum presenting a linear theoretical analysis made of the performance of three proportional navigation guidance systems installed in a given supersonic, variable-incidence, boost-glide, antiaircraft missile at Mach numbers of 2.7 and 1.3. Three guidance systems are compared on the basis of the maximum obtainable speed of response of the missile and guidance-system combination consistent with adequate stability. Results regarding the effect of method of positioning the radar antenna on the speed of response, effect of Mach number on response, and effect of networks are provided.

Report presenting the low-speed longitudinal characteristics of swept wings derived from investigations at high Reynolds numbers. Two different types of flow separation, trialing-edge and leading-edge separation, are identified and discussed.

Report presenting a theoretical investigation of the longitudinal dynamic characteristics of an automatically stabilized supersonic canard missile configuration with a target seeker sensitive to changes in pitch altitude. Variables explored include seeker gain, time delay, and non-linearities, including dead spots in the seeker.

Report presenting results of pressure-distribution measurements at low speed on a wing-fuselage combination with a wing of W plan form with aspect ratio 6, 45 degrees sweep, taper ratio 0.6, and an NACA 65A009 airfoil section placed parallel to the plane of symmetry. Results regarding aerodynamic characteristics, pressure distribution on the wing, pressure distribution on the fuselage, aerodynamic section characteristics, and stall patterns are provided.

Report presenting an experimental investigation to determine the thrust augmentation possible by the injection of liquid ammonia into the compressor inlet of an axial-flow-type turbojet engine. Results regarding the selection of coolant, cooling with liquid ammonia, engine performance, and operating experience are provided.

Report presenting an investigation of large-chord flap-type controls with sweptback hinge lines and with various sizes and shapes of horn balances on a triangular wing of aspect ratio 2. Changes in the following elements were investigated: shielding of the horn balance, percentage of horn balance, trailing-edge thickness, flap nose seal, contour of the horn balance, and Reynolds number.

Report presenting an investigation to determine the importance of molecular-scale processes in the overall turbojet combustion process. The effect of oxygen concentration on the inlet oxygen-nitrogen mixture on the combustion efficiency of a single J33 combustor was determined for a range of combustor-inlet pressures and a range of fuel-flow rates. At a given fuel-flow rate, combustion efficiency decreased at an increasing rate with reduction in oxygen concentration.

Report presenting measurements in free fall at transonic speeds of the dynamic stability characteristics of three models. Two of the models had 45 degree swept wings of aspect ratio 6 and 45 degree swept tail surfaces and differed only in that one had plane wings and one was cambered and twisted. The third had the same fuselage-tail arrangement but no wing.

"An investigation of the jet-spreading characteristics of a 16 inch ram-jet engine was conducted in the 8 by 6 foot supersonic tunnel at a Mach number of 2.0; both a converging nozzle having a contraction ratio of 0.71 and a cylindrical extension to the combustion chamber were used. The jet boundaries determined by means of pitot pressure surveys were compared with boundaries calculated from one-dimensional continuity and momentum relations. For the cylindrical nozzle, the jet reaches its maximum diameter, 4 percent greater than calculated, about 0.6 nozzle-exit diameter downstream of the nozzle exit" (p. 1).

Report presenting the results of an investigation to determine the effects of simulated skin wrinkling of the wing surface on the aerodynamic characteristics of two wing-body combinations. The lift, drag, and pitching moment of a low-wing wing-body combination with a plane triangular wing of aspect ratio 3 are presented for a range of Mach and Reynolds numbers.

Memorandum presenting minimum spark-ignition engines at reduced pressures for ethane, ethylene, acetylene, n-hexane, cyclohexane, and benzene.

Report presenting rocket-propelled models flown at transonic and supersonic speeds to determine the effect on zero-lift drag of increasing the leading-edge bluntness of a sweptback wing. Results regarding the total drag coefficient, wing-plus-interference drag coefficients, and force-break Mach number are provided.

Report presenting a flight investigation using four rocket-propelled models of low-aspect-ratio tailless configurations between Mach numbers 0.7 and 1.35. The configurations investigated were a 45 degree delta model with NACA 63A010 airfoil sections parallel to free stream, a 45 degree delta-wing model with NACA 63A006 airfoil sections parallel to free stream, a 37.5 degree swept-tapered wing model with NACA 64(sub 1)A012 sections perpendicular to quarter chord, and a 39 degree swept-tapered-wing model with NACA 64A006 airfoil sections perpendicular to quarter chord. Results regarding time histories, static longitudinal stability, dynamic longitudinal stability, order of oscillation frequency, rotational damping in pitch, and drag are provided.

Report presenting an investigation to determine the effects of turbine-blade coatings, supplied by fuel additives, on heat transfer with combustor exhaust gases to an S-816 alloy blade. Two fuel additives were used. The results indicated that the fuel additives provided an oxide coating on the combustor-liner wall and other metal parts, including the blade, exposed to the exhaust gas.