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  Partner: UNT Libraries Government Documents Department
 Serial/Series Title: NACA Advanced Confidential Report
 Collection: National Advisory Committee for Aeronautics Collection
The Supersonic Axial-Flow Compressor
An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained. digital.library.unt.edu/ark:/67531/metadc60311/
Extension of Useful Operating Range of Axial-Flow Compressors by Use of Adjustable Stator Blades
A theory has been developed for resetting the blade angles of an axial-flow compressor in order to improve the performance at speeds and flows other than the design and thus extend the useful operating range of the compressor. The theory is readily applicable to the resetting of both rotor and stator blades or to the resetting of only the stator blades and is based on adjustment of the blade angles to obtain lift coefficients at which the blades will operate efficiently. Calculations were made for resetting the stator blades of the NACA eight-stage axial-flow compressor for 75 percent of design speed and a series of load coefficients ranging from 0.28 to 0.70 with rotor blades left at the design setting. The NACA compressor was investigated with three different blade settings: (1) the design blade setting, (2) the stator blades reset for 75 percent of design speed and a load coefficient of 0.48, and (3) the stator blades reset for 75 percent of design speed and a load coefficient of 0.65. digital.library.unt.edu/ark:/67531/metadc60225/
Correlation of Wright Aeronautical Corporation cooling data on the R-3350-14 intermediate engine and comparison with data from the Langley 16-foot high-speed tunnel
No Description digital.library.unt.edu/ark:/67531/metadc62698/
Paths of target-seeking missiles in two dimensions
No Description digital.library.unt.edu/ark:/67531/metadc62682/
Paths of Target Seeking Missiles in Two Dimensions
Parameters that enter into equation of trajectory of a missile are discussed. Investigation is made of normal pursuit, of constant, proportional, and line--of-sight methods of navigation employing target seeker, and of deriving corresponding pursuit paths. Pursuit paths obtained under similar conditions for different methods are compared. Proportional navigation is concluded to be best method for using target seeker installed in missile. digital.library.unt.edu/ark:/67531/metadc65596/
Parameters determining performance of supersonic pilotless airplanes powered by ram-compression power planes
No Description digital.library.unt.edu/ark:/67531/metadc62640/
Investigations on laminar boundary-layer stability and transition on curved boundaries
No Description digital.library.unt.edu/ark:/67531/metadc62701/
Stability and control force tests of four- and six-unit wing designs of low aspect ratio and 20 degree triangular plan form
No Description digital.library.unt.edu/ark:/67531/metadc62680/
Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps
Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils. digital.library.unt.edu/ark:/67531/metadc61295/
Field of Flow About a Jet and Effect of Jets on Stability of Jet-Propelled Airplanes
A theoretical investigation was conducted on jet-induced flow deviation. Analysis is given of flow inclination induced outside cold and hot jets and jet deflection caused by angle of attack. Applications to computation of effects of jet on longitudinal stability and trim are explained. Effect of jet temperature on flow inclination was found small when thrust coefficient is used as criterion for similitude. The average jet-induced downwash over tail plane was obtained geometrically. digital.library.unt.edu/ark:/67531/metadc62555/
Approximate formulas for the computation of turbulent boundary-layer momentum thicknesses in compressible flows
No Description digital.library.unt.edu/ark:/67531/metadc62569/
Development of Wing Inlets
Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil. digital.library.unt.edu/ark:/67531/metadc61340/
Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint
Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025. digital.library.unt.edu/ark:/67531/metadc61323/
Flight investigation at high speeds of profile drag of wing of a P-47D airplane having production surfaces covered with camouflage paint
No Description digital.library.unt.edu/ark:/67531/metadc53369/
A method for the calculation of external lift, moment, and pressure drag of slender open-nose bodies of revolution at supersonic speeds
No Description digital.library.unt.edu/ark:/67531/metadc62664/
Theory and application of hot-wire instruments in the investigation of turbulent boundary layers
No Description digital.library.unt.edu/ark:/67531/metadc62409/
Determination of the stability and control characteristics of a straight-wing, tailless fighter-airplane model in the Langley free-flight tunnel
No Description digital.library.unt.edu/ark:/67531/metadc61003/
Effects of specific types of surface roughness on boundary-layer transition
No Description digital.library.unt.edu/ark:/67531/metadc61337/
Effects of specific types of surface roughness on boundary-layer transition
No Description digital.library.unt.edu/ark:/67531/metadc53334/
Occurrence of iron oxides on cast-iron engine surfaces after operation
No Description digital.library.unt.edu/ark:/67531/metadc61910/
Effect of Compressibility on the Pressure and Forces Acting on a Modified NACA 65,3-019 Airfoil Having a 0.20-Chord Flap
An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value. digital.library.unt.edu/ark:/67531/metadc61397/
The effect of wall interference upon the aerodynamic characteristics of an airfoil spanning a closed-throat circular wind tunnel
No Description digital.library.unt.edu/ark:/67531/metadc61348/
Wind-tunnel investigation of an NACA 66-series 16-percent-thick low-drag tapered wing with Fowler and split flaps
No Description digital.library.unt.edu/ark:/67531/metadc61410/
A flight investigation of NACA aileron modifications for the improvement of the lateral control characteristics of a high-speed fighter airplane
No Description digital.library.unt.edu/ark:/67531/metadc61647/
Two-dimensional wind-tunnel investigation of 0.20-airfoil-chord plain ailerons of different contour on an NACA 65(sub 1)-210 airfoil section
No Description digital.library.unt.edu/ark:/67531/metadc61394/
Effects of Compressibility on the Maximum Lift Characteristics and Spanwise Load Distribution of a 12-Foot-Span Fighter-Type Wing of NACA 230-Series Airfoil Sections
Lift characteristics and pressure distribution for a NACA 230 wing were investigated for an angle of attack range of from -10 to +24 degrees and Mach range of from 0.2 to 0.7. Maximum lift coefficient increased up to a Mach number of 0.3, decreased rapidly to a Mach number of 0.55, and then decreased moderately. At high speeds, maximum lift coefficient was reached at from 10 to 12 degrees beyond the stalling angle. In high-speed stalls, resultant load underwent a moderate shift outward. digital.library.unt.edu/ark:/67531/metadc61423/
An infrared cloud indicator I : analysis of infrared-radiation exchange with tables and chart for calibration of the cloud indicator
No Description digital.library.unt.edu/ark:/67531/metadc62395/
Rubber conductors for aircraft ignition cables
No Description digital.library.unt.edu/ark:/67531/metadc62514/
Comparison of fixed-stabilizer, adjustable- stabilizer and all-movable horizontal tails
No Description digital.library.unt.edu/ark:/67531/metadc61649/
An Experimental Investigation of NACA Submerged-Duct Entrances
The results of a preliminary investigation of submerged duct entrances are presented. It is shown that an entrance of this type possess desirable critical speed and pressure recovery characteristics when used on a fuselage or nacelle in a region of low incremental velocity and thin boundary layer. The data obtained indicate that submerged entrances are most suitable for use with internal-flow systems which diffuse the air only a small amount: for example, those used with jet motors which have axial-flow compressors. Where complete diffusion of the air is required, fuselage-nose or wing leading edge inlets may prove to be superior. The results of the investigation have been prepared in such a form as to permit their use by a designer and the application of these data to a specific design is discussed. digital.library.unt.edu/ark:/67531/metadc64879/
Flight investigation of boundary-layer and profile-drag characteristics of smooth wing sections of a P-47D airplane
No Description digital.library.unt.edu/ark:/67531/metadc53374/
Knock-limited power outputs from a CFR engine using internal coolants II : six aliphatic amines
No Description digital.library.unt.edu/ark:/67531/metadc62055/
Supersonic-tunnel tests of projectiles in Germany and Italy
No Description digital.library.unt.edu/ark:/67531/metadc62434/
Bibliography and review of information relating to the hydrodynamics of seaplanes
No Description digital.library.unt.edu/ark:/67531/metadc62558/
Effect of wing modifications on the longitudinal stability of a tailless all-wing airplane model
No Description digital.library.unt.edu/ark:/67531/metadc61001/
Charts for helicopter-performance estimation
No Description digital.library.unt.edu/ark:/67531/metadc61830/
A simple method for estimating terminal velocity including effect of compressibility on drag
No Description digital.library.unt.edu/ark:/67531/metadc61252/
Blade design data for axial-flow fans and compressors
No Description digital.library.unt.edu/ark:/67531/metadc62634/
The development and application of high-critical-speed nose inlets
No Description digital.library.unt.edu/ark:/67531/metadc279653/
Effect of leakage past aileron nose on aerodynamic characteristics of plain and internally balanced ailerons on NACA 66(215)-216, a = 1.0 airfoil
No Description digital.library.unt.edu/ark:/67531/metadc61639/
Effect of the lift coefficient on propeller flutter
No Description digital.library.unt.edu/ark:/67531/metadc62153/
An additional investigation of the high-speed lateral-control characteristics of spoilers
No Description digital.library.unt.edu/ark:/67531/metadc60955/
Completed Tabulation in the United States of Tests of 24 Airfoils at High Mach Numbers (Derived from Interrupted Work at Guidonia, Italy in the 1.31- by 1.74-Foot High-Speed Tunnel)
Two-dimensional data were obtained in Mach range of from 0.40 to 0.94 and Reynolds Number range of (3.4 - 4.2) X 10 Degrees. Results indicate that thickness ratio is dominating shape parameter at high Mach numbers and that aerodynamic advantages are attainable by using thinnest possible sections. Effects of jet boundaries, Reynolds Number, and Data presented are free from jet-boundary and humidity effects. digital.library.unt.edu/ark:/67531/metadc61347/
The Effect of Inlet Pressure and Temperature on the Efficiency of a Single Stage Impulse Turbine Having an 11.0-Inch Pitch-Line Diameter Wheel
Efficiency tests have been conducted on a single-stage impulse engine having an 11-inch pitch-line diameter wheel with inserted buckets and a fabricated nozzle diaphragm. The tests were made to determine the effect of inlet pressure, Inlet temperature, speed, and pressure ratio on the turbine efficiency. An analysis is presented that relates the effect of inlet pressure and temperature to the Reynolds number of the flow. The agreement between the analysis and the experimental data indicates that the changes in turbine efficiency with Inlet pressure and temperature may be principally a Reynolds number effect. digital.library.unt.edu/ark:/67531/metadc61859/
Effects on low-speed spray characteristics of various modifications to a powered model of the Boeing XPBB-1 flying boat
No Description digital.library.unt.edu/ark:/67531/metadc61764/
Performance of compressor-turbine jet-propulsion systems
No Description digital.library.unt.edu/ark:/67531/metadc62464/
A correlation of the effects of compression ratio and inlet-air temperature on the knock limits of aviation fuels in a CFR engine I
No Description digital.library.unt.edu/ark:/67531/metadc62295/
Flight investigation of the variation of drag coefficient with Mach number for the Bell P-39N-1 airplane
No Description digital.library.unt.edu/ark:/67531/metadc61128/
Flight tests of dive-recovery flaps on an XP-51 airplane
No Description digital.library.unt.edu/ark:/67531/metadc61739/
Preliminary Investigation of Supersonic Diffusers
No Description digital.library.unt.edu/ark:/67531/metadc62636/
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