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At least 289 records · Page 16

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NASA NTRS1932

Compression-ignition engine tests of several fuels

The tests reported in this paper were made to devise simple engine tests which would rate fuels as to their comparative value and their suitability for the operating conditions of the individual engine on which the tests are made. Three commercial fuels were used in two test engines having combustion chambers with and without effective air flow. Strictly comparative performance tests gave almost identical results for the three fuels. Analysis of indicator cards allowed a differentiation between fuels on a basis of rates of combustion. The same comparative ratings were obtained by determining the consistent operating range of injection advance angle for the three fuels. The difference in fuels is more pronounced in a quiescent combustion chamber than in one with high-velocity air flow. A fuel is considered suitable for the operating conditions of an engine with a quiescent combustion chamber if it permits the injection of the fuel to be advanced beyond the optimum without exceeding allowable knock or allowable maximum cylinder pressures.

Spanogle, J A↗

NASA NTRS1987

Engine flow visualization using a copper vapor laser

A flow visualization system has been developed to determine the air flow within the combustion chamber of a motored, axisymmetric engine. The engine has been equipped with a transparent quartz cylinder, allowing complete optical access to the chamber. A 40-Watt copper vapor laser is used as the light source. Its beam is focused down to a sheet approximately 1 mm thick. The light plane is passed through the combustion chamber, and illuminates oil particles which were entrained in the intake air. The light scattered off of the particles is recorded by a high speed rotating prism movie camera. A movie is then made showing the air flow within the combustion chamber for an entire four-stroke engine cycle. The system is synchronized so that a pulse generated by the camera triggers the laser's thyratron. The camera is run at 5,000 frames per second; the trigger drives one laser pulse per frame. This paper describes the optics used in the flow visualization system, the synchronization circuit, and presents results obtained from the movie. This is believed to be the first published study showing a planar observation of airflow in a four-stroke piston-cylinder assembly. These flow visualization results have been used to interpret flow velocity measurements previously obtained with a laser Doppler velocimetry system.

Regan, Carolyn A.↗

NASA NTRS1990

Numerical analysis of bipropellant combustion in orbit maneuvering vehicle thrust chamber

The combustion and aerodynamic processes of liquid-liquid bipropellant; monomethylhydrazine and nitrogen tetroxide in the combustion chamber of a variable thrust engine (VTE) have been examined to assess the flow structure, combustion characteristics, and wall heat transfer under the simulated engine operations at two power levels. The Generalized Elliptic Multi-phase flow with Chemical reaction, Heat transfer and Interfacial Processes (GEM-CHIP) code and the Bipropellant Combustion (BICOMB) code, which is the GEMCHIP algorithm reformulated in the frame work of a body fitted coordinate system, have been used for the numerical study. Combustion flow structure with non-premixed type injection that simulates the impingement type injector is predicted and the results are compared with that of the premixed injector.

Chiu, H. H.↗

NASA NTRS1998

Method and Apparatus for Coupling Space Vehicles

A first space vehicle is attached with a second space vehicle which includes a rocket propulsion nozzle having a combustion chamber upstream from the nozzle throat. Apparatus 10 includes an elongate grappling arm 12 extending from the first space vehicle, a pair of inflatable bladders 16, 18 positioned about the grappling arm for engaging an inner surface of the combustion chamber upon inflation. and a pair of rear bladders 20, 22 positioned about the grappling arm for engaging an inner surface of the nozzle downstream from the combustion chamber upon inflation for aligning the grappling arm and the rocket propulsion nozzle. A pressurized fluid source 24 is provided on the first space vehicle for supplying fluid pressure to the inflatable bladders. and a fluid control valve manifold 30 selectively controls the release of pressurized fluid to the bladders. According to the method of the invention. the grappling arm is inserted into the rocket propulsion nozzle. and the control valves are actuated to first inflate the front bladders and thereby interconnect the grappling arm and the rocket propulsion nozzle. The rear bladders are subsequently inflated to align a central axis of the grappling arm with a central axis of the rocket propulsion nozzle. Inflation of the rear bladders provides an axial reaction load to balance the axial load provided by the front bladders. Attaching the vehicles in space may be controlled from the earth by activating the control valves to inflate the bladders. The cost of attaching space vehicles is significantly reduced by attaching a grappling arm on one vehicle with the existing rocket propulsion nozzle of another vehicle.

Schneider, William C.↗

NASA NTRS2018

Development of New Rocket Engine Liner Materials Based on Copper Alloys with Diamond Particle Additions

Novel copper-diamond materials containing particulate dispersions of diamond within a copper alloy matrix have been produced as next-generation rocket engine combustion chamber liner materials. These Copper-Diamond (Cu-D) composite materials have significantly higher thermal conductivity than conventional liner materials and have the potential to increase the engine performance significantly. Liner fabrication techniques include hot pressing and spark plasma sintering (SPS), which are capable of scale up to full-size production parts. The materials have been evaluated for thermal cycling resistance and ease of manufacturing. The results indicate great promise of these materials and associated economics to replace state-of-the-art combustion chamber liner materials such as GRCop-84 and NARloy-Z. Combustion chamber fabrication processes are also discussed, including SPS and additive manufacturing.

Bhat, Biliyar N.↗

NASA NTRS2018

Development of New Rocket Engine Liner Materials Based on Copper Alloys with Diamond Particle Additions

Novel copper-diamond materials containing particulate dispersions of diamond within a copper alloy matrix have been produced as next-generation rocket engine combustion chamber liner materials. These Copper-Diamond composite materials have significantly higher thermal conductivity than conventional liner materials and have the potential to increase the engine performance significantly. Liner fabrication techniques include hot pressing and spark plasma sintering, which are capable of scale up to full-size production parts. The materials have been evaluated for thermal cycling resistance and ease of manufacturing. The results indicate great promise of these materials and associated economics to replace state-of-the-art combustion chamber liner materials such as GRCop-84 and NARloy-Z. Novel combustion chamber fabrication processes are discussed, including cladding the liner with jacket materials.

Bhat, Biliyar N.↗

NASA NTRS2000

Catalyst Bed Instability Within the USFE H2O2/JP-8 Rocket Engine

Orbital Sciences Corporation has been awarded a contract by NASA's Marshall Space Flight Center, in cooperation with the U.S. Air Force Research Laboratory's Military Space Plane Technology Program Office, for the Upper Stage Flight Experiment (USFE) program. Orbital is designing, developing, and will flight test a new low-cost, 10,000 lbf hydrogen peroxide/ JP-8 pressure fed liquid rocket. During combustion chamber tests at NASA Stennis Space Center (SSC) of the USFE engine, the catalyst bed showed a low frequency instability occurring as the H202 flow reached about 1/3 its design rate. This paper reviews the USFE catalyst bed and combustion chamber and its operation, then discusses the dynamics of the instability. Next the paper describes the dynamic computer model used to recreate the instability. The model was correlated to the SSC test data, and used to investigate possible solutions to the problem. The combustion chamber configuration which solved the instability is shown, and the subsequent stable operation presented.

Johnson, Curtis W.↗

NASA NTRS1999

Technology and Advanced Development for a Non-Toxic Orbital Maneuvering System and Reaction Control System for Orbiter Upgrade

NASA intends to pursue technology applications to upgrade the Space Shuttle Orbiter OMS and RCS systems with non-toxic propellants. The primary objectives of an upgraded OMS/RCS are improved safety and reliability, reduced operations and maintenance costs while meeting basic OMS/RCS operational and performance requirements. The OMS/RCS has a high degree of direct interaction with the crew and requires subsystem and components that are compatible with integration into the orbiter vehicle with regard to external mold-line, power and thermal control The non-toxic propulsion technology is also applicable to future Human Exploration and Development of Space (HEDS) missions. The HEDS missions have similar requirements for attitude control and lander descent/ascent propulsion and which will emphasize the use of In-Situ Resource for propellants. When used as a regenerative coolant as in the Shuttle Orbiter OMS combustion chamber, non-toxic fuels such as ethanol are limited in their cooling capacity by the bulk temperature rise permitted to prevent film boiling or possible coking. Typical regeneratively cooled chambers are constructed from highly conductive copper, which maximizes heat transfer, or from low conductivity materials like stainless steel that can also exacerbate cooling problems. For an ethanol cooled application the heat transfer into the fluid must be controlled to reduce the fuel coolant bulk temperature rise. An approach to provide this control is the subject of this report. This report is being issued to document work done by Aerojet on NASA contract NAS 8-98042. Specifically, this project investigates of the use of ethanol, a designated non-toxic fuel, as a coolant for the Space Shuttle Orbital Maneuvering System Engine combustion chamber. The project also addresses a cost reducing fabrication technique for construction of such a combustion chamber. The study contained three major sub-tasks: an analytical investigation and trade study which included layout of a flight type chamber concept, the fabrication and evaluation of formed platelet liner panels and the preparation and testing of mechanical properties specimens representative of a novel hot gas wall concept.

Hayes, W. A.↗

NASA NTRS2011

Plasma Igniter for Reliable Ignition of Combustion in Rocket Engines

A plasma igniter has been developed for initiating combustion in liquid-propellant rocket engines. The device propels a hot, dense plasma jet, consisting of elemental fluorine and fluorine compounds, into the combustion chamber to ignite the cold propellant mixture. The igniter consists of two coaxial, cylindrical electrodes with a cylindrical bar of solid Teflon plastic in the region between them. The outer electrode is a metal (stainless steel) tube; the inner electrode is a metal pin (mild steel, stainless steel, tungsten, or thoriated-tungsten). The Teflon bar fits snugly between the two electrodes and provides electrical insulation between them. The Teflon bar may have either a flat surface, or a concave, conical surface at the open, down-stream end of the igniter (the igniter face). The igniter would be mounted on the combustion chamber of the rocket engine, either on the injector-plate at the upstream side of the engine, or on the sidewalls of the chamber. It also might sit behind a valve that would be opened just prior to ignition, and closed just after, in order to prevent the Teflon from melting due to heating from the combustion chamber.

Martin, Adam↗

NASA NTRSDate not supplied

Development of Large-Scale Thrust Chambers using Laser Powder Directed Energy Deposition

Historical manufacturing methods for liquid propulsion combustion chambers have relied on series of tube-walls or slotted channel walls (subsequently brazed or plated) for regenerative cooling paths. Large-scale additive manufacturing techniques offer strong potential to minimize manufacturing lead times and cost for developing liquid propulsion chambers, while maintaining performance metrics of traditionally manufactured hardware. The benefits expand beyond traditional combustion chambers to lesser considered nuclear thermal propulsion chambers where radiation energizes the singular incoming propellant stream instead of an exothermic chemical reaction. While each subset of liquid propulsion chambers reaps similar benefit, the geometric challenges in each design can vary substantially. This manuscript discusses the challenges of large contraction ratio propulsion chambers, bimetallic liners and closeout jackets, and material selection when considering laser powder directed energy deposition. LP-DED presently offers build volumes unattainable by L-PBF systems, required for large diameter thrust chambers. LP-DED typically has faster deposition time and is capable of overhang angles without support material (due to the ability to tilt the build plate and minimize the actual build angle). These benefits come at the cost of feature resolution. Designers and manufacturers should consider the effects of thick-wall features printed parallel to the primary surface versus perpendicular due to distortion. Thick-wall and thin-wall features printed in the same plane can substantially increase build time. Where large-scale L-PBF overlaps in size, there may be limited benefit in the quantity of powder required for the build due to the low capture rate of LP-DED and a case-by-case examination is required.

LP-DED↗

NASA NTRS2001

Low-Cost Gas Generator and Ignitor

A gas generator and ignitor assembly for powering a turbine of a turbopump assembly for a rocket engine comprises an injector and a combustion chamber. the injector having a body member including a fuel inlet and an oxidizer inlet spaced one from the other and communicating with respective radially spaced apart annular members in the body member. Three annuli communicate with the fuel inlet and two annuli communicate with the oxidizer inlet. the annuli which communicates with the oxidizer being positioned between pairs of the other annuli, The body member is enclosed by a plate having an array of bores arranged in two series with three radially spaced apart groups of circular rows in each series. The outer series has 28 groups of triplet bores while the inner series has 14 groups of triplet bores. The annuli which communicate with oxidizer feed bores of each series that are between the other bores of a triplet. the latter bores communicating with annuli that communicate with fuel. The inner and outer bores of the triplets of each series are inclined relatively to each other and to the third bore of the triplet so that fuel and oxidizer atomizes as it is sprayed into the inlet of the combustion chamber where the propellants are mixed. and burned. The burning of the propellants is effected by ignition of a plug of solid propellant fuel mounted to communicate with the interior of the combustion chamber.

Henry J Dennis, Jr↗

NASA NTRS1995

Creep Testing of High-Temperature Cu-8 Cr-4 Nb Alloy Completed

A Cu-8 at.% Cr-4 at.% Nb (Cu-8 Cr-4 Nb) alloy is under development for high-temperature, high heatflux applications, such as actively cooled, hypersonic vehicle heat exchangers and rocket engine combustion chambers. Cu-8 Cr-4 Nb offers a superior combination of strength and conductivity. It has also shown exceptional low-cycle fatigue properties. Following preliminary testing to determine the best processing route, a more detailed testing program was initiated to determine the creep lives and creep rates of Cu-8 Cr-4 Nb alloy specimens produced by extrusion. Testing was conducted at the NASA Lewis Research Center with constant-load vacuum creep units. Considering expected operating temperatures and mission lives, we developed a test matrix to accurately determine the creep properties of Cu-8 Cr-4 Nb between 500 and 800 C. Six bars of Cu-8 Cr-4 Nb were extruded. From these bars, 54 creep samples were machined and tested. The figure on the left shows the steady-state, or second-stage, creep rates for the samples. Comparison data for NARloy-Z (Cu-3 wt % Ag-0.5 wt % Zr), the alloy currently used in combustion chamber liners, were not unavailable. Therefore the steady-state creep rates for Cu at similar temperatures are presented. As expected, in comparison to pure Cu, the creep rates for Cu-8 Cr-4 Nb are much lower. The lives of the samples are presented in the figure on the right. As shown, Cu-8 Cr-4 Nb at 800 C is comparable to NARloy-Z at 648 C. At equivalent temperatures, Cu-8 Cr-4 Nb enjoys a 20 to 50 percent advantage in stress for a given life and 1 to 3 orders of magnitude greater life at a given stress. The improved properties allow for design tradeoffs and improvements in new and existing heat exchangers such as the next generation of combustion chamber liners. Average creep rates for Cu-8 Cr-4 Nb and pure Cu are shown. Average creep lives for Cu-8 Cr- 4 Nb and NARloy-Z are also shown. Currently, two companies are interested in the commercial usage of the Cu-8 Cr-4 Nb alloy. The Rocketdyne Division of Rockwell International is conducting independent testing to analyze the properties for their projected needs in advanced rocket engine applications. Metallamics, a company based in Traverse City, Michigan, is entering into a Space Act Agreement to evaluate and test Cu-Cr-Nb alloys as materials for welding electrodes that are used in robotic welding operations. Creep rate is one of the alloy properties that determines the degree to which a welding electrode will mushroom or expand at the tip. A material with a low creep rate will resist mushrooming and give the electrode a longer life, minimizing downtime. This application holds the potential for large-scale usage of the alloy in the automotive and other industries. Success here would dramatically decrease the cost of the alloy and increase availability for aerospace applications.

Source record↗

NASA NTRS1990

Combustor design and analysis using the Rocket Combustor Interactive Design (ROCCID) methodology

The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure.

Klem, Mark D.↗

NASA NTRS1990

Combustor design and analysis using the ROCket Combustor Interactive Design (ROCCID) Methodology

Klem, Mark D.↗

NASA NTRS1987

Hydrogen scramjet with sidewall injection - Shock tunnel simulations

A free-piston shock tunnel has been used to obtain test data on a scramjet combustion chamber with sidewall injection. The results obtained indicate that combustion was strongly influenced by a region of fuel whose temperature was held below its ignition temperature by wall-cooling effects; this increased the fraction of unburned fuel and resulted in a significant loss of specific impulse. Aerodynamic heating would keep the walls above hydrogen ignition temperature in an actual scramjet powerplant, however. Maximum specific impulse was obtained with a combination of parallel and transverse injection in a long combustion chamber, followed by a dual stage expansion.

Morgan, R. G.↗

NASA NTRS1989

Shock-Induced Heating In A Rocket Engine

Misalignments give rise to hotspots on walls. Report discusses numerical simulation of flow in and near small, ringlike cavity in wall of Space Shuttle main engine at junction of main combustion chamber and nozzle. Purpose to study effects of misalignments between combustion chamber and nozzle on transfer of heat into surfaces chamber, cavity, and nozzle. Depending on specific misalignment flow encounters forward-or backward-facing step leaving chamber and entering nozzle. Results in serious losses of performance and excessive heating of walls.

Lagnado, Ronald R.↗

NASA NTRS1988

Finite area combustor theoretical rocket performance

Previous to this report, the computer program of NASA SP-273 and NASA TM-86885 was capable of calculating theoretical rocket performance based only on the assumption of an infinite area combustion chamber (IAC). An option was added to this program which now also permits the calculation of rocket performance based on the assumption of a finite area combustion chamber (FAC). In the FAC model, the combustion process in the cylindrical chamber is assumed to be adiabatic, but nonisentropic. This results in a stagnation pressure drop from the injector face to the end of the chamber and a lower calculated performance for the FAC model than the IAC model.

Gordon, Sanford↗

NASA NTRS1996

Numerical Modeling of Spray Combustion with an Unstructured-Grid Method

The present unstructured-grid method follows strictly the basic finite volume forms of the conservation laws of the governing equations for the entire flow domain. High-order spatially accurate formulation has been employed for the numerical solutions of the Navier-Stokes equations. A two-equation k-epsilon turbulence model is also incorporated in the unstructured-grid solver. The convergence of the resulted linear algebraic equation is accelerated with preconditioned Conjugate Gradient method. A statistical spray combustion model has been incorporated into the present unstructured-grid solver. In this model, spray is represented by discrete particles, rather than by continuous distributions. A finite number of computational particles are used to predict a sample of total population of particles. Particle trajectories are integrated using their momentum and motion equations and particles exchange mass, momentum and energy with the gas within the computational cell in which they are located. The interaction calculations are performed simultaneously and eliminate global iteration for the two-phase momentum exchange. A transient spray flame in a high pressure combustion chamber is predicted and then the solution of liquid-fuel combusting flow with a rotating cup atomizer is presented and compared with the experimental data. The major conclusion of this investigation is that the unstructured-grid method can be employed to study very complicated flow fields of turbulent spray combustion. Grid adaptation can be easily achieved in any flow domain such as droplet evaporation and combustion zone. Future applications of the present model can be found in the full three-dimensional study of flow fields of gas turbine and liquid propulsion engine combustion chambers with multi-injectors.

Shang, H. M.↗