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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 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

A Review Towards the Design Optimization of High-Performance Additively Manufactured Rotating Detonation Rocket Engine Injectors

Rotating Detonation Rocket Engines (RDRE) have been marketed primarily for their higher specific impulse potential over constant pressure (CP) liquid rocket engines. However, several other performance advantages exist such as heat transfer advantages for gas expander cycle, increased completeness of combustion at low chamber L*, compact engine design, reduced coolant channel pressure drop potential, and improved injector C* performance. NASA has paved the way for liquid engine system performance enhancement since the Apollo program and continues to do so with metal additive manufacturing (AM), new advanced materials, and advanced propulsion concepts. A team of propulsion development engineers at NASA are in the process of developing high-performance 7K lbf class RDRE hardware for their potential use in lander, upper stage, and even launch vehicle applications. Clear advantages have been demonstrated with AM including program cost and schedule reductions of up to 50%. It is well known that injector performance is integrally linked to the global performance of a combustion device. This is especially the case for RDREs since detonation stability is heavily dependent on the mixedness of propellants. A major goal of this work is to identify what has been done in the open experimental literature and what injectors design features are conducive to high performance in the detonation cycle. This paper reviews the available literature and reports the primary gaps in the knowledge base needed by the pressure gain combustion (PGC) community. Major conclusions are documented, and suggestions given towards the design of high-performance liquid RDRE injectors. In addition, the integration of metal AM into the design of liquid RDRE injector schemes is included.

rotating detonation rocket engine↗
NASA NTRS1963

MODELING TECHNIQUES FOR LIQUID PROPELLANT ROCKET COMBUSTION PROCESSES

A qualitative physical processes description of stable and unstable combustion in rockets is presented to establish the basis for choosing model hardware design criteria. It is concluded that it is not possible to scale rocket combustion processes in the usual sense of the term. The studies conducted at Rocketdyne have shown that model chambers which satisfactorily model the steady state behavior of large engines must be designed to maintain (1) the propellant injection density, and (2) the chamber to throat Contraction ratio (hence chamber pressure). For studies of destructive acoustic modes of combustion instability a third parameter, the frequency, must also be maintained.

COMBUSTION↗
NASA NTRS1964

The Effects of Wind on Sounding Rockets

A theory to account for the effects of wind on sounding rockets is developed and discussed. A rapid method for calculating the effects of wind is presented. The results of a series of calculations are compared with experimental values. Several suggestions are made to reduce the sensitivity of sounding rockets to winus and to provide better predictions.

SOUNDING ROCKET↗
NASA NTRS1964

Effect of Thrustor Arcing on Ion Rocket System Design

Arcing between the electrodes of an ion rocket thrustor is investigated for the effect it will have on the overall design of ion rocket systems. The effect of generator, transformer, and inserted impedances on system stability and efficiency, and on operating lifetime of thrustor components are considered. Rate of arcing, preferential arc paths, speed of arc interruption, and response of thrustors to voltage transients are also discussed. The study indicates that arcing is a system hazard and that arcing rates are of prime importance in the design and development of thrustors, thrustor arrays, and system components. Two system models for reducing the hazard are proposed, based on principles developed herein, published data, and experience available from operation of nine-module electron-bombardment thrustor array. Experimental investigation of these models is presently underway.

THRUSTOR↗
NASA NTRS1965

Investigation of light hydrocarbon fuels with flox mixtures as liquid rocket propellants final report, 30 jun. 1964 - 30 jun. 1965

An analysis was completed for determination of the most promising hydrocarbon fuels for use with flox in upper stage rocket engines. Experimental rocket firings in uncooled, transpiration cooled, and regeneratively cooled thrust chambers were conducted using flox with methane, propane, butene-1, and a eutectic blend of pentane and isopentane. Experimental heated tube heat transfer and hypergolicity tests were conducted, and laboratory determinations of physical properties of various blends of hydrocarbon compounds were made.

HYDROCARBON FUEL↗
NASA NTRS1956

Flight Investigation of the Performance of a Two-stage Solid-propellant Nike-deacon (DAN) Meteorological Sounding Rocket

A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

MISSILES, SPECIFIC TYPES↗
NASA NTRS1953

An Investigation of High-frequency Combustion Oscillations in Liquid-propellant Rocket Engines

An experimental investigation of high-frequency combustion oscillations (screaming) was conducted with a 100-pound-thrust acid-hydrocarbon rocket engine and a 500-pound-thrust oxygen-fuel rocket engine. The oscillation frequencies could be correlated as a linear function of the parameter C/L, where C is the experimentally measured characteristic velocity and L is the combustion-chamber length. The tendency of the engines to scream increased as chamber length was increased. With engine configurations that normally had a low efficiency, screaming resulted in increased performance; at the same time, a five to tenfold increase in heat-transfer rate occurred. It was possible, however, to achieve good performance without screaming.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1994

SSTO rockets. A practical possibility

Most experts agree that single-stage-to-orbit (SSTO) rockets would become feasible if more advanced technologies were available to reduce the vehicle dry weight, increase propulsion system performance, or both. However, these technologies are usually judged to be very ambitious and very far off. This notion persists despite major advances in technology and vehicle design in the past decade. There appears to be four major misperceptions about SSTOs, regarding their mass fraction, their presumed inadequate performance margin, their supposedly small payloads, and their extreme sensitivity to unanticipated vehicle weight growth. These misperceptions can be dispelled for SSTO rockets using advanced technologies that could be matured and demonstrated in the near term. These include a graphite-composite primary structure, graphite-composite and Al-Li propellant tanks with integral reusable thermal protection, long-life tripropellant or LOX-hydrogen engines, and several technologies related to operational effectiveness, including vehicle health monitoring, autonomous avionics/flight control, and operable launch and ground handling systems.

PROPELLANT MASS FRACTION↗
NASA NTRS2014

Developing Avionics Hardware and Software for Rocket Engine Testing

My summer was spent working as an intern at Kennedy Space Center in the Propulsion Avionics Branch of the NASA Engineering Directorate Avionics Division. The work that I was involved with was part of Rocket University's Project Neo, a small scale liquid rocket engine test bed. I began by learning about the layout of Neo in order to more fully understand what was required of me. I then developed software in LabView to gather and scale data from two flowmeters and integrated that code into the main control software. Next, I developed more LabView code to control an igniter circuit and integrated that into the main software, as well. Throughout the internship, I performed work that mechanics and technicians would do in order to maintain and assemble the engine.

liquid rocket↗
NASA NTRS2017

Methods for Reducing Singly Reflected Rays on the Wolter-I Focusing Figures of the FOXSI Rocket Experiment

In high energy solar astrophysics, imaging hard X-rays by direct focusing offers higher dynamic range and greater sensitivity compared to past techniques that used indirect imaging. The Focusing Optics X-ray Solar Imager (FOXSI) is a sounding rocket payload which uses seven sets of nested Wolter-I figured mirrors that, together with seven high-sensitivity semiconductor detectors, observes the Sun in hard X-rays by direct focusing. The FOXSI rocket has successfully flown twice and is funded to fly a third time in Summer 2018. The Wolter-I geometry consists of two consecutive mirrors, one paraboloid, and one hyperboloid, that reflect photons at grazing angles. Correctly focused X-rays reflect twice, once per mirror segment. For extended sources, like the Sun, off-axis photons at certain incident angles can reflect on only one mirror and still reach the focal plane, generating a pattern of single-bounce photons that can limit the sensitivity of the observation of faint focused X-rays. Understanding and cutting down the singly reflected rays on the FOXSI optics will maximize the instrument's sensitivity of the faintest solar sources for future flights. We present an analysis of the FOXSI singly reflected rays based on ray-tracing simulations, as well as the effectiveness of different physical strategies to reduce them.

FOXSI Rocket Experiment↗
NASA NTRS2017

Application of Background Oriented Schlieren for Altitude Testing of Rocket Engines

A series of experiments was performed to determine the feasibility of using the Background Oriented Schlieren, BOS, flow visualization technique to image a simulated, small, rocket engine, plume under altitude test conditions. Testing was performed at the NASA Glenn Research Centers Altitude Combustion Stand, ACS, using nitrogen as the exhaust gas simulant. Due to limited optical access to the facility test capsule, all of the hardware required to conduct the BOS were located inside the vacuum chamber. During the test series 26 runs were performed using two different nozzle configurations with pressures in the test capsule around 0.3 psia. No problems were encountered during the test series resulting from the optical hardware being located in the test capsule and acceptable resolution images were captured. The test campaign demonstrated the ability of using the BOS technique for small, rocket engine, plume flow visualization during altitude testing.

Background Oriented Schlieren↗
NASA NTRS2020

First Operation of TES Microcalorimeters in Space with the Micro-X Sounding Rocket

Micro-X is a sounding rocket-borne instrument that uses a microcalorimeter array to perform high-resolution X-ray spectroscopy. This instrument flew for the first time on July 22nd, 2018, from the White Sands Missile Range, USA. This flight marks the first successful operation of a transition-edge sensor array and its time-division multiplexing readout system in space. This launch was dedicated to the observation of the supernova remnant Cassiopeia A. A failure in the attitude control system prevented the rocket from pointing and led to no time on target. The onboard calibration source provided X-rays in flight, and it is used to compare detector performance during preflight integration, flight, and after the successful post-flight recovery. This calibration data demonstrate the capabilities of the detector in a space environment as well as its potential for future flight.

TES↗