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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 559 records · Page 31
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Al6061-RAM2 Development and Hot-Fire Testing using Additive Manufacturing Laser Powder Directed Energy Deposition for Liquid Rocket Engine Channel-Cooled Nozzles

Aluminum 6061-RAM2 is a high-strength aluminum feedstock developed for additive manufacturing (AM)processes. This alloy leverages Reactive Additive Manufacturing (RAM) technology. The RAM aluminum alloys were developed to be weldable—therefore printable—while equaling or exceeding strength properties of high strength wrought aluminum alloys. NASA and industry partners developed Laser Powder Directed Energy Deposition (LP-DED)additive manufacturing of Al6061-RAM2 for use in aerospace applications. Efforts included establishing build parameters, characterizing the alloy, fabricating components, and completing hot-fire testing of complex internal channel-cooled nozzles. These efforts are to address the growing need for large-scale parts using high-performance light-weight materials. Two rocket engine nozzles were fabricated using LP-DED Al6061-RAM2 that included integral cooling channels. The Al6061-RAM2 has completed process development and initial properties were established. This paper provides an overview of the LP-DED process development, material characterization and properties, component manufacturing, supplemental development, and hot-fire testing. Results from hot-fire testing are provided for a lander-class 31kN (7,000lbf)thrust engine using Liquid Oxygen (LOX)/Liquid Hydrogen(LH2) and LOX/Methane(LCH4).

Al6061-RAM2↗
NASA NTRSDate not supplied

Al6061-RAM2 Development and Hot-Fire Testing using Additive Manufacturing Laser Powder Directed Energy Deposition for Liquid Rocket Engine Channel-Cooled Nozzles

Aluminum 6061-RAM2 is a high-strength aluminum feedstock developed for additive manufacturing (AM)processes. This alloy leverages Reactive Additive Manufacturing (RAM) technology. The RAM aluminum alloys were developed to be weldable—therefore printable—while equaling or exceeding strength properties of high strength wrought aluminum alloys. NASA and industry partners developed Laser Powder Directed Energy Deposition (LP-DED)additive manufacturing of Al6061-RAM2 for use in aerospace applications. Efforts included establishing build parameters, characterizing the alloy, fabricating components, and completing hot-fire testing of complex internal channel-cooled nozzles. These efforts are to address the growing need for large-scale parts using high-performance light-weight materials. Two rocket engine nozzles were fabricated using LP-DED Al6061-RAM2 that included integral cooling channels. The Al6061-RAM2 has completed process development and initial properties were established. This paper provides an overview of the LP-DED process development, material characterization and properties, component manufacturing, supplemental development, and hot-fire testing. Results from hot-fire testing are provided for a lander-class 31kN (7,000lbf) thrust engine using Liquid Oxygen (LOX)/Liquid Hydrogen (LH2) and LOX/Methane (LCH4).

Al6061-RAM2↗
NASA NTRSDate not supplied

Scoping of an Air Supply Configuration for AFC on a Commercial Transport Airplane High Lift System

A scoping study is being conducted on the air supply for a microjet-based active flow control (AFC) system on a twin-turbofan commercial transport airplane. Microjets provide circulation control using small surface-normal pneumatic jets located near the trailing edge of a lifting surface such as a wing or flap. When located on the pressure side of the lifting surface they increase the lift, and when located on the suction side they decrease lift. In this study, microjets are considered for installation in the flaps of the high-lift version of the Common Research Model (CRM-HL). An air supply system involving bleed air from the airplane’s auxiliary power unit (APU) plus ram air provides the air to the microjets. A model based on the 1D compressible flow equations is applied to analyze the air supply system configuration and predict the microjet flow rate with the resulting airplane performance changes based on Reynolds-averaged Navier-Stokes modeling of microjets on the CRM-HL. The results of this scoping study are encouraging in that APU air can be used to entrain ram air and thereby increase the AFC mass flow rate to achieve effective lift control and airplane performance enhancement during takeoff and landing.

AATT↗
NASA NTRS1952

Pressure Recovery, Drag, and Subcritical Stability Characteristics of Three Conical Supersonic Diffusers at Stream Mach Numbers from 1.7 to 2.0

A study of a 20 degree and a 25 degree half-angle high mass-flow ratio conical supersonic inlet was made on a 16-inch ram jet in the 8- by 6-foot supersonic tunnel. A greater range of stable subcritical operation was obtained with the low mass-flow ratio inlets; a greater range was obtained with the 25 degree than with the 20 degree half-angle low mass-flow ratio inlet. The high mass-flow ratio inlet had the least drag.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1952

Spreading of Exhaust Jet from 16 Inch Ream Jet at Mach Number 2.0 / Fred Wilcox, Donald Pennington

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. The maximum diameter for the converging nozzle was 7 percent greater than calculated from one dimensional relations and occurred from 1 to 1.5 nozzle-exit diameters downstream of the exit. Non dimensional maximum jet diameters agreed closely with results of an investigation by Rousso and Baughman; these data were obtained with low-temperature jets exhausting into a stream at a Mach number of 1.91 from nozzles having exit diameters of 0.75 inch.

INTERFERENCE, JET-MISSILES↗
NASA NTRS1964

Upsetting Butt Edge Increases Weld-Joint Strength

Mechanical upsetting /a mode of cold forging/ of butt edges to be welded is accomplished by the use of hydraulic rams and pressure rollers. The mechanical upsetting increases the thickness of the material in the heat-affected zone and compensates for the lower specific strength per unit thickness common to this area.

Vesco, D.↗
NASA NTRS1956

Effect of Engine Operating Variables on Combustor-inlet Parameters

Studies of the fundamental processes of combustion are usually concerned with wide ranges of investigation of individual processes. In general, each fundamental combustion process may be studied in an environment that is most suited to its evaluation and possibly unrelated basically to any practical application. The majority of the data presented in volume I of this series concern the fundamental aspects of combustion as functions of the individual occurrence of various contributing processes. In a jet engine, however, the various fundamental combustion processes may occur simultaneously and may interact. Furthermore, the engine environment usually does not permit independent variation of single combustion parameters, since specified operating conditions impose specific values on the parameters. In volume II, data are presented to show the effect of operating conditions on the over-all combustion process in different combustion components. To show the effect of operating conditions, it is necessary to specify the range of these conditions within which combustion components may operate. Therefore, this chapter presents only the operating conditions that might be required in the primary combustors and afterburners of typical current turbojet engines. (Corresponding information on ram-jet engines is presented in ch. xisi.) This chapter is not intended to serve as an explanation of engine operation. The operating conditions of the combustion components are presented in terms of total pressures and temperatures at the primary-combustor and afterburner inlets, reference velocities and outlet total temperatures of the primary combustors, and velocities at the plane of the flameholder in the afterburners. The data are presented to relate the operating regions of typical current turbojet combustion components to flight altitudes, Mach numbers, and modes of engine operation. Specifically, data are presented for the combustion parameters of the primary combustor and afterburner of three turbojet engines having rated compressor total-pressure ratios of 5, 8, and 12 under full-throttle conditions. Operational data for the primary combustor also include part-throttle operation at 70, 80, and 90 percent of rated engine speed and windmifling operation. The range of flight conditions includes altitudes from sea level to 65,000 feet and flight Mach numbers from zero to 1.6.

Jones, A. W.↗
NASA NTRS1968

Fuel transfer system permits rapid coupling

Docking and fuel transfer system provides an efficient method for transferring fuel from a tanker to another vehicle. With this system, no triggering operation is required prior to docking, the support system can be rigidized by simply locking the rams of shock absorbers, and no separate fuel line coupling action is required.

West, A. M.↗
NASA NTRS1969

Development of an Advanced Annular Combustor

The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

Rusnak, J. P.↗
NASA NTRS1970

Reliability Analysis Model

RAM program determines probability of success for one or more given objectives in any complex system. Program includes failure mode and effects, criticality and reliability analyses, and some aspects of operations, safety, flight technology, systems design engineering, and configuration analyses.

Source record↗
NASA NTRS1971

Electrostatic Probe Measurements of Plasma Surrounding Three 25,000 Foot Per Second Reentry Flight Experiments

A high-altitude electrostatic probe experiment is described and data are presented for three RAM C reentries at velocities of approximately 25 OOO feet per second. The electron density profiles inferred from the probe measurements are compared with theoretical calculations, and the probable cause of significant disagreement is briefly discussed. Probe measurements during material-addition sequences for radio-blackout alleviation are presented, and comparisons of probe measurements with other pertinent diagnostic measurements are made.

W Linwood Jones, Jr↗
NASA NTRS1971

Subsonic and transonic pressure distributions around a bluff afterbody in the wake of a 120 deg cone for various separation distances

A wind-tunnel investigation was conducted at free-stream Mach numbers from 0.20 to 1.00 and Reynolds numbers, based on maximum afterbody diameter, from 2.25 million to 6.90 million on solid models of an attached inflatable decelerator (AID) concept. Tests were conducted to obtain static and ram surface pressure distributions about the basic shapes and at various separation distances between the 120 deg conical forebody and the inflated afterbody shape. The resulting data were used to study the feasibility of extracting a payload from a conical forebody by means of an AID.

Sawyer, J. W.↗
NASA NTRS1972

Thermal design study of an air-cooled plug-nozzle system for a supersonic cruise aircraft

A heat-transfer design analysis has been made of an air-cooled plug-nozzle system for a supersonic-cruise aircraft engine. The proposed 10deg half-angle conical plug is sting supported from the turbine frame. Plug cooling is accomplished by convection and film cooling. The flight profile studied includes maximum afterburning from takeoff to Mach 2.7 and supersonic cruise at Mach 2.7 with a low afterburner setting. The calculations indicate that, for maximum afterburning, about 2 percent of the engine primary flow, removed after the second stage of the nine-stage compressor, will adequately cool the plug and sting support. Ram air may be used for cooling during supersonic-cruise operations, however. Therefore, the cycle efficiency penalty paid for air cooling the plug and sting support should be low.

Clark, J. S.↗
NASA NTRS1972

Subsonic pressure distributions around a solid model of an inflatable decelerator attached to the base of an ogive-cylinder

A wind-tunnel investigation was conducted at free-stream Mach numbers from 0.20 to 1.00 and corresponding Reynolds numbers, based on maximum afterbody diameter, from 2.25 x one million to 6.90 x one million on a solid model of an attached inflatable decelerator (AID) connected to the base of an ogive-cylinder. Tests were conducted to obtain ram-air and surface pressure distributions about the AID. AID shapes derived for subsonic deployment are dependent on the pressure distributions used in their derivation, and the different shapes obtained are dependent on the Mach number for which the design is made. The resulting pressure distributions were used in a design program to obtain new shapes which were compared with the original pressure-distribution shape.

Sawyer, J. W.↗