Report regarding experiments made to determine the fallout-radiation protection offered by five typical residential structures located at the Nevada Test Site.

Memorandum presenting flight tests conducted with a director-type radar fire-control system, with scope presentation of the steering information, to determine the effects of attack-display smoothing on the pilot's tracking and steering effectiveness in tail-chase and lead-collision beam attacks. Results regarding tail-chase attacks and lead-collision beam attacks are provided.

Memorandum presenting a study of aerodynamic performance and static stability and control at hypersonic speeds. In the first part of the study, the effect of interference lift is investigated by tests of asymmetric models with conical fuselages and arrow plan-form wings. In the second part, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail.

Note presenting a simplified method for determination of the critical height of three-dimensional roughness particles required to promote premature transition of a laminar boundary layer on models of airplanes or airplane components in a wind tunnel with zero heat transfer. The method has been applied to various types of configurations in several wind-tunnel investigations.

Memorandum presenting the effect of blunting on the internal performance of a two-cone inlet with an internally cylindrical cowl investigated experimentally at Mach number 4.95. Total-pressure-recovery and mass-flow data were obtained for a range of spike and cowl bluntness.

Report presenting an investigation of the effect of inlet component blunting on performance using an axisymmetric external-compression inlet in the supersonic wind tunnel. The investigation was carried out to determine the performance penalties associated with spike-tip and cowl-lip blunting. Results regarding the effect of rounding the spike tip on inlet performance, overall drag coefficients, flow distortion, and Schileren photographs are provided.

Report discussing an investigation of the dynamic stability and control characteristics of a model of the Ryan X-13 jet VTOL airplane in hovering and transition flight. Flight control was smooth and easy without the gyroscopic effects of the engine represented, but adding a simulation of the gyroscopic effects caused the plane to become uncontrollable in hovering flight without artificial stabilization. The transition was found to be easy to perform with and without gyroscopic effects, but the model had a tendency to fly in roll and slideslipped attitudes.

Report discussing a study of the effects of puncturing the heat shield of an intercontinental-ballistic-missile warhead by small projectiles. Calculations were created for both rod and sphere projectiles and experimental testing was performed on a missile model with holes drilled in the heat shield. The possibility that a projectile could have enough energy to cause mechanical damage to the interior of the warhead is also presented.

"Characteristic exhaust velocity was measured for a small range of mixture ratios with four different injectors. Performances of parallel-sheet, like-on-like, and triplet injectors were about the same, but a parallel-jet injector had a much lower performance. Performance values for ammonia-oxygen were slightly lower than for heptane-oxygen" (p. 1).

"A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor" (p. 1).

"Investigation of the control parameters of an external-internal compression inlet indicates that the cowl-lip shock provides a signal to position the spike and to start the inlet over a Mach number range from 2.1 to 3.0. Use of a single fixed probe position to control the spike over the range of conditions resulted in a 3.7-count loss in total-pressure recovery at Mach 3.0 and 0 deg angle of attack. Three separate shock-sensing-probe positions were required to set the spike for peak recovery from Mach 2.1 to 3.0 and angles of attack from 0 deg to 6 deg" (p. 1).

Memorandum presenting an experimental investigation of an internal-contraction, axially symmetric inlet with isentropic compression surfaces on both the cowl and centerbody conducted over a range of Mach numbers from 2.0 to 2.7 at angles of attack from 0 to 6 degrees. The study was made to determine the optimum bleed system, together with a method of controlling the inlet.

Memorandum presenting an investigation in an altitude test chamber to evaluate the use of highly reactive chemicals injected into a turbojet afterburner to promote the combustion process, which was inhibited by water vapor from compressor-inlet injection. The chemicals evaluated were commercial hydrogen and aluminum trimethyl. Results regarding the effects of hydrogen injection on afterburner performance and effect of aluminum trimethyl afterburner fuel additive on performance are provided.

Report presenting an investigation of the off-design performance of fixed- and variable-geometry divergent ejectors. Results regarding fixed geometry and low secondary flow, fixed geometry and high secondary flow, variable geometry and low secondary flow, variable geometry and high secondary flow, and ejectors with full afterburning are provided.

Skin temperature and surface pressure of blunted cone-cylinder-flare configuration free flight test vehicle to hypersonic speeds.

"Heat-transfer measurements were made on a simulated glide-rocket shape in free flight at Mach numbers up to 10 and free-stream Reynolds numbers of 2 x 10 based on distance along surface from apex and 3 x 10 based on nominal leading-edge diameter. The model simulated the bottom of a 75 deg delta wing at 80 deg angle of attack. The data indicated that for the test conditions a modified three-dimensional stagnation-point theory will predict to reasonable engineering accuracy the heating on a highly swept wing leading edge, the heating being reduced by sweep by the 3/2 power of the cosine of the sweep angle" (p. 1).

Results regarding a wind-tunnel investigation of wingless missile configurations with cylindrical bodies and conical or hemispherical noses, extensible control surfaces aft of the nose, and tails consisting of eight low-aspect-ratio triangular or rectangular fins. Normal-force, axial-force, and pitching-moment coefficients were obtained for several control deflections for Mach numbers 1.2 and 1.9.

Memorandum presenting a design and experimental investigation of a single-stage supersonic turbine. The turbine was designed for a rotor entering relative Mach number of 2. Results regarding the overall turbine performance, outer-wall static-pressure variation, momentum-loss considerations, effect of rotor modifications, and supersonic starting are provided.

Memorandum presenting an investigation to determine the feasibility of a high-performance transonic axial-flow compressor stage with a weight flow of 40 pounds per second per square foot of frontal area. A transonic axial-flow inlet stage with a hub-tip ratio of 0.35 and an axial Mach number of approximately 0.75 was designed and fabricated. Results regarding overall rotor performance, flow parameters, radial matching of blade-element sections, and comparison of blade-element parameters with design rules are provided.

"The buffeting loads acting on the wing and horizontal tail of a 1/4-scale model of the X-1E airplane have been measured in the Langley 16-foot transonic tunnel in the Mach number range from 0.40 to 0.90. When the buffeting loads were reduced to a nondimensional aerodynamic coefficient of buffeting intensity, it was found that the maximum buffeting intensity of the horizontal tail was about twice as large as that of the wing. Comparison of power spectra of buffeting loads acting on the horizontal tail of the airplane and of the model indicated that the model horizontal tail, which was of conventional force-test-model design, responded in an entirely different mode than did the airplane" (p. 1).

Turbulence was generated by using screens, and the turbulence percentage was measured by a hot-wire anemometer both in the boundary layer and the free stream. The local heat-transfer coefficient was measured at 12 locations along the plate for the cases of various turbulence levels. The transition Reynolds number from laminar to turbulent flow decreases as the main-stream turbulence level increases. In the range of laminar heat transfer the effect of turbulence in the main flow was not great, but in the range of turbulent heat transfer the heat-transfer coefficient increases according to the increase of turbulence.

"A number of semiempirical approximate methods exist for determining the characteristics of the turbulent boundary layer on a curvilinear surface. At present, among these methods, the one proposed by L. G. Loitsianskii is given frequent practical application. This method is sufficiently effective and permits, in the case of wing profiles with technically smooth surfaces, calculating the basic characteristics of the boundary layer and the values of the overall drag with an accuracy which suffices for practical purposes" (p. 1).

Analysis is presented on the possible similarity solutions of the three-dimensional, laminar, incompressible, boundary-layer equations referred to orthogonal, curvilinear coordinate systems. Requirements of the existence of similarity solutions are obtained for the following: flow over developable surface and flow over non-developable surfaces with proportional mainstream velocity components.

The method of coordinate perturbation is applied to the unsteady flow of a compressible fluid in ducts of variable cross section. Solutions, in the form of perturbation series, are obtained for unsteady flows in ducts for which the logarithmic derivative of area variation with respect to the space coordinate is a function of the 'smallness' parameter of the perturbation series. This technique is applied to the problem of the interaction of a disturbance and a shock wave in a diffuser flow. It is found that, for a special choice of the function describing the disturbance, the path of the shock wave can be expressed in closed form to first order. The method is then applied to the determination of the flow field behind a shock wave moving on a prescribed path in the x,t-plane. Perturbation series solutions for quite general paths are developed. The perturbation series solutions are compared with the more exact solutions obtained by the application of the method of characteristics. The approximate solutions are shown to be in reasonably accurate agreement with the solutions obtained by the method of characteristics.