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Analytical investigation of the effect of water injection on supersonic turbojet-engine - inlet matching and thrust augmentation

Description: From Introduction: "In order to provide engine-inlet matching to prevent such mass flow spillage and the attendant drags, much mechanical complication has been introduced. References 1 and 2 propose employing this principle at subsonic and supersonic speeds to achieve thrust augmentation. The merits of the system for supersonic engine-inlet airflow matching, as well as thrust augmentation, appear considerable; and a comparative study with existing fixed- and variable-geometry inlets was therefore undertaken."
Date: January 1957
Creator: Beke, Andrew
Partner: UNT Libraries Government Documents Department

Experimental Investigation of Water Injection in Subsonic Diffuser of a Conical Inlet Operation at Free-Stream Mach Number of 2.5

Description: A spike-type nose inlet with sharp-lip cowl was investigated at a free-stream Mach number of 2.5 with water injection in its 16-inch diameter, 11-foot-long subsonic diffuser section. Inlet total temperature of exit with liquid-air ratios of about 0.04 with no apparent change in the critical pressure recovery. The observed temperature drops were less than the theoretically predicted values, and the amount of water evaporated was 35 to 50 percent less than that theoretically possible.
Date: January 15, 1957
Creator: Beke, Andrew
Partner: UNT Libraries Government Documents Department

Effect of Wall Cooling on Inlet Parameters of a Scoop Operating in a Turbulent Boundary Layer on a Flat or Conical Surface for Mach Numbers 2 to 10

Description: Note presenting analytical results obtained for boundary-layer mass flow, momentum, total-temperature, and total-pressure recovery ratios of a scoop inlet with a height equal to the boundary-layer thickness and operating in a turbulent boundary layer, for flat and conical surfaces with wall cooling. The results indicate that design of boundary-layer intake devices or diverters must reflect the effects of wall cooling by either variable or compromised size.
Date: March 1958
Creator: Beke, Andrew
Partner: UNT Libraries Government Documents Department

Thrust and drag characteristics of simulated variable-shroud nozzles with hot and cold primary flows at subsonic and supersonic speeds

Description: Report presenting an experimental investigation of a series of fixed-geometry exhaust nozzles which simulated a variable exit, which was conducted in the 8- by 6-foot supersonic wind tunnel at Mach numbers 0, 0.6, 1.5, 1.7, and 2.0. Gross thrust performance of the long-shroud exit configurations was considerably better than the short-shroud. Results regarding flow coefficients, thrust characteristics, flight performance analysis of data, and shroud drag are provided.
Date: February 18, 1955
Creator: Beke, Andrew & Simon, Paul C.
Partner: UNT Libraries Government Documents Department

Performance comparison at supersonic speeds of inlets spilling excess flow by means of bow shock, conical shock, or bypass

Description: Report presenting a comparison of fixed-geometry, translating-spike, and bypass-inlets on the basis of turbojet- and ramjet-engine performance. Results regarding a comparison of the experimental data and its application to ramjet and turbojet engines are provided.
Date: October 23, 1953
Creator: Allen, J. L. & Beke, Andrew
Partner: UNT Libraries Government Documents Department

Force and pressure recovery characteristics at supersonic speeds of a conical spike inlet with bypasses discharging in an axial direction

Description: Report presenting an investigation of an axially symmetric nacelle-type conical spike inlet with two bypasses located in the horizontal plane and on opposite sides of the nacelle in the 8- by 6-foot supersonic tunnel at several Mach numbers and angles of attack. Results regarding the performance with open bypasses and with closed bypasses are provided.
Date: January 30, 1953
Creator: Allen, J. L. & Beke, Andrew
Partner: UNT Libraries Government Documents Department

Force and pressure-recovery characteristics of a conical-type nose inlet operating at Mach numbers of 1.6 to 2.0 and at angles of attack to 9 degrees

Description: Report presenting an investigation of an axially symmetric spike-type nose inlet suitable for a nacelle power-plant installation in the 8- by 6-foot supersonic tunnel at a range of Mach numbers and angles of attack. The inlet was designed to attain a mass-flow ratio of unity a Mach number 2.0. Results regarding the performance of inlet A and inlet B, which differed based on subsonic diffuser area variation, are provided.
Date: November 19, 1952
Creator: Beke, Andrew & Allen, J. L.
Partner: UNT Libraries Government Documents Department

Force and pressure recovery characteristics at supersonic speeds of a conical spike inlet with a bypass discharging from the top or bottom of the diffuser in an axial direction

Description: Force and pressure-recovery characteristics of a nacelle-type conical-spike inlet with a fixed-area bypass located in the top or bottom of the diffuser are presented for flight Mach numbers of 1.6, 1.8, and 2.0 for angles of attack from 0 degrees to 9 degrees. Top or bottom location of the bypass did not have significant effects on diffuser pressure-recovery, bypass mass-flow ratio, or drag coefficient over the range of angles of attack, flight Mach numbers, and stable engine mass-flow ratios investigated. A larger stable subcritical operating range was obtained with the bypass on the bottom at angles of attack from 3 degrees to 9 degrees at a flight Mach number of 2.0. At a flight Mach number of 2.0, the discharge of 14 percent of the critical mass flow of the inlet by means of a bypass increased the drag only one-fifth of the additive drag that would result for equivalent spillage behind an inlet normal shock without significant reductions in diffuser pressure recovery.
Date: March 23, 1953
Creator: Allen, J. L. & Beke, Andrew
Partner: UNT Libraries Government Documents Department

Performance Characteristics of a Normal-Shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8

Description: The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems, angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface.
Date: July 29, 1952
Creator: Dryer, Murray & Beke, Andrew
Partner: UNT Libraries Government Documents Department

Effect of an Adjustable Supersonic Inlet on the Performance Up to Mach Number 2.0 of a J34 Turbojet Engine

Description: "A J34 turbojet engine was investigated at free-stream Mach numbers of 0.12 and 1.6 to 2.0 to determine the effect of supersonic inlet operation on engine performance. With the exception of ideal jet thrust, the use of generalized engine parameters correlated the engine data satisfactorily when the exit nozzle was choked. Large total-pressure distortions did not affect compressor efficiency" (p. 1).
Date: January 4, 1956
Creator: Beke, Andrew; Englert, Gerald & Beheim, Milton
Partner: UNT Libraries Government Documents Department

Preliminary investigation of a conical spike inlet in combination with a vertical-wedge auxiliary inlet at Mach number 1.9

Description: Pressure-recovery characteristics of a nacelle-type-spike inlet in combination with a vertical-wedge auxiliary scoop are presented for a free-stream Mach number of 1.9 at zero angle of attack. The auxiliary scoop provided 17 percent additional air flow with a drop in critical pressure recovery from 0.86 to 0.81. However, in terms of inlet-engine matching, the pressure recovery of the undersized spike inlet operating at a specified corrected air flow increased with the scoop open, for example, from 0.69 to 0.81.
Date: September 27, 1955
Creator: Beke, Andrew; Allen, John L. & Williams, Thomas
Partner: UNT Libraries Government Documents Department

Thrust characteristics of a series of convergent-divergent exhaust nozzles at subsonic and supersonic flight speeds

Description: From Summary: "An experimental investigation of a series of four convergent-divergent exhaust nozzles was conducted in the Lewis 8-by-6 foot supersonic wind tunnel at Mach numbers of 0.1, 0.6, 1.6, and 2.0 over a range of nozzle pressure ratios. The thrust characteristics of these nozzles were determined by a pressure-integration technique. From a thrust standpoint, a nozzle designed to give uniform parallel flow at the exit had no advantage over the simple geometric design with conical convergent and divergent sections."
Date: March 12, 1954
Creator: Fradenburgh, Evan A.; Gorton, Gerald C. & Beke, Andrew
Partner: UNT Libraries Government Documents Department