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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 181 records · Page 10

Fabrication of Composite Combustion Chamber/Nozzle for Fastrac Engine

The Fastrac Engine developed by the Marshall Space Flight Center for the X-34 vehicle began as a low cost engine development program for a small booster system. One of the key components to reducing the engine cost was the development of an inexpensive combustion chamber/nozzle. Fabrication of a regeneratively cooled thrust chamber and nozzle was considered too expensive and time consuming. In looking for an alternate design concept, the Space Shuttle's Reusable Solid Rocket Motor Project provided an extensive background with ablative composite materials in a combustion environment. An integral combustion chamber/nozzle was designed and fabricated with a silica/phenolic ablative liner and a carbon/epoxy structural overwrap. This paper describes the fabrication process and developmental hurdles overcome for the Fastrac engine one-piece composite combustion chamber/nozzle.

Lawerence, T.↗

Fabrication of Composite Combustion Chamber/Nozzle for Fastrac Engine

The Fastrac Engine developed by the Marshall Space Flight Center for the X-34 vehicle began as a low cost engine development program for a small booster system. One of the key components to reducing the engine cost was the development of an inexpensive combustion chamber/nozzle. Fabrication of a regeneratively cooled thrust chamber and nozzle was considered too expensive and time consuming. In looking for an alternate design concept, the Space Shuttle's Reusable Solid Rocket Motor Project provided an extensive background with ablative composite materials in a combustion environment. An integral combustion chamber/nozzle was designed and fabricated with a silica/phenolic ablative liner and a carbon/epoxy structural overwrap. This paper describes the fabrication process and developmental hurdles overcome for the Fastrac engine one-piece composite combustion chamber/nozzle.

Lawrence, T.↗

Effect of silicone oil on solid propellant combustion in small motors

The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.

Ramohalli, K.↗

Electroslag welding of 75 to 100 ton ingots

A new process for joining large cross-section components by electroslag welding is reported that utilizes four consumable electrodes made of the same material as the ingots. A bifilar current supply provides the very large energy input to the consumerable nozzle for heating the ingots being welded. This process produces better and more homogeneous mechanical properties, better chemical composition and structure uniformity, and requires no prior or simultaneous heating of ingots during welding. It is now possible to obtain medium size ingots and join as many as are required to produce 500 or 600 ton ingots.

Manning, C. R., Jr.↗

Quiet Clean Short-haul Experimental Engine (QCSEE). Under-The-Wing (UTW) engine boilerplate nacelle test report. Volume 3: Mechanical performance

Results of initial tests of the under the wing experimental engine and boilerplate nacelle are presented. The mechanical performance of the engine is reported with emphasis on the advanced technology components. Technology elements of the propulsion system covered include: system dynamics, composite fan blades, reduction gear, lube and accessory drive system, fan frame, inlet, core cowl cooling, fan exhaust nozzle, and digital control system.

Source record↗

Computer-Aided Process Model For Carbon/Phenolic Materials

Computer program implements thermochemical model of processing of carbon-fiber/phenolic-matrix composite materials into molded parts of various sizes and shapes. Directed toward improving fabrication of rocket-engine-nozzle parts, also used to optimize fabrication of other structural components, and material-property parameters changed to apply to other materials. Reduces costs by reducing amount of laboratory trial and error needed to optimize curing processes and to predict properties of cured parts.

Letson, Mischell A.↗

Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

Fisher, Mark F.↗

Fabrication and Testing of Ceramic Matrix Composite Propulsion Components

A viewgraph presentation outlines NASA's goals for the Second and Third Generation Reusable Launch Vehicles, placing emphasis on improving safety and decreasing the cost of transporting payloads to orbit. The use of ceramic matrix composite (CMC) technology is discussed. The development of CMC components, such as the Simplex CMC Blisk, cooled CMC nozzle ramps, cooled CMC thrust chambers, and CMC gas generators, are described, including challenges, test results, and likely future developments.

Effinger, Michael R.↗

Response Surface Modeling of Combined-Cycle Propulsion Components using Computational Fluid Dynamics

Three examples of response surface modeling with CFD are presented for combined cycle propulsion components. The examples include a mixed-compression-inlet during hypersonic flight, a hydrogen-fueled scramjet combustor during hypersonic flight, and a ducted-rocket nozzle during all-rocket flight. Three different experimental strategies were examined, including full factorial, fractionated central-composite, and D-optimal with embedded Plackett-Burman designs. The response variables have been confined to integral data extracted from multidimensional CFD results. Careful attention to uncertainty assessment and modeling bias has been addressed. The importance of automating experimental setup and effectively communicating statistical results are emphasized.

Steffen, C. J., Jr.↗

Flight Results and Lessons Learned from the Starling Propulsion System

NASA’s Starling mission is a swarm of four 6U Cubesats flying in formation in Low Earth Orbit to demonstrate scalable swarm technologies, particularly related to crosslink networking. In order to maintain and modify their formation, the Starling spacecraft use a cold gas propulsion system called Hamlet, which provides each spacecraft with 30 m/s of delta-V, as well as attitude control during maneuvers. Hamlet is a fully self-contained system, incorporating propellant tanks, valves, nozzles, and control electronics. Most of the structure of Hamlet is a single piece of stereolithography-printed composite, which simplifies assembly and allows for unusual tank geometry that maximizes propellant volume in the allocated space. An extensive qualification campaign was conducted for Hamlet, including performance characterization that has been largely validated by in-flight measurements. Since the start of nominal mission operations in late summer 2023, Hamlet has conducted weekly or bi-weekly maneuvers of up to 0.29 m/s to assemble the swarm, maintain formation, and change swarm configurations as required by the experiments. Operations have not been without challenges, including a propellant leak on one unit and thrust variability issues on all four. This paper describes the design and implementation of Hamlet, as well as in-flight performance data, anomaly investigation and resolution, and lessons learned.

Starling↗

Advanced Materials for the Lunar Surface: Multiscale Computational Design of Refractory Alloys and Carbides

Emerging operational environments, such as the lunar surface, present novel challenges for NASA and drive the need for advanced materials in applications like fission surface power systems. To address these demands, computational materials science is rapidly evolving to augment or replace costly and hazardous empirical testing. Although materials selection at NASA remains predominantly experimentally driven, advanced simulation methodologies are being steadily integrated into the engineering lifecycle. This work details the application of multiscale simulation techniques—including first-principles calculations, CALPHAD, dislocation dynamics, and molecular dynamics—at NASA's Ames Research Center to evaluate advanced materials for extreme environments. First, we present contributions to the Space Nuclear Propulsion Project. Be-cause propellant channel coatings in nuclear thermal rockets must withstand high-pressure, high-temperature hydro-gen, optimizing these materials is critical. First-principles calculations were employed to establish a rigorous quantitative and qualitative understanding of the behavior of the refractory carbides ZrC, NbC, and their mixtures in high-enthalpy hydrogen environments. This necessitated the generation of high-fidelity thermodynamic models for both stoichiometric and carbon-depleted carbides, both with and without the presence of hydrogen. Furthermore, we highlight efforts under the Refractory Alloy Additive Manufacturing Build Optimization (RAAMBO) project, where existing and novel alloy compositions were assessed for additive manufacturing printability and subsequent performance in applications such as heat pipes and rocket nozzle extensions. This was accomplished through a comprehensive multiscale simulation framework that bridged the gap from the nanometer to the millimeter scale. Across both initiatives, rigorous validation against empirical data was prioritized. By systematically employing a verified and validated computational frame-work, we demonstrate how simulation effectively supports multidisciplinary engineering efforts, builds project-wide confidence, and drives critical materials development.

computational materials↗

Further Improvements to Nozzle Boundary Layer Calculations in BLIMPJ

Further improvements made to advance the current Boundary Layer Integral Matrix Procedure - Version J (BLIMPJ) containing previously modeled simplified calculation methods by accounting for condensed phase, thick boundary layer and free stream turbulence effects are discussed. The condensed phase effects were included through species composition effect considered via input to the code and through particle damping effect considered via a turbulence model. The thrust loss calculation procedure for thick boundary layer effects was improved and the optimization of net thrust with respect to nozzle length was performed. The effects of free stream turbulence were approximately modeled in the turbulence model.

Praharaj, S. C.↗

Quiet Clean Short-haul Experimental Engine (QCSEE) Under-The-Wing (UTW) composite nacelle

The detail design of the under the wing experimental composite nacelle components is summarized. Analysis of an inlet, fan bypass duct doors, core cowl doors, and variable fan nozzle are given. The required technology to meet propulsion system performance, weight, and operational characteristics is discussed. The materials, design, and fabrication technology for quiet propulsion systems which will yield installed thrust to weight ratios greater than 3.5 to 1 are described.

Johnston, E. A.↗

High-Melt Carbon-Carbon Coating for Nozzle Extensions

Carbon-Carbon Advanced Technologies, Inc. (C-CAT), has developed a high-melt coating for use in nozzle extensions in next-generation spacecraft. The coating is composed primarily of carbon-carbon, a carbon-fiber and carbon-matrix composite material that has gained a spaceworthy reputation due to its ability to withstand ultrahigh temperatures. C-CAT's high-melt coating embeds hafnium carbide (HfC) and zirconium diboride (ZrB2) within the outer layers of a carbon-carbon structure. The coating demonstrated enhanced high-temperature durability and suffered no erosion during a test in NASA's Arc Jet Complex. (Test parameters: stagnation heat flux=198 BTD/sq ft-sec; pressure=.265 atm; temperature=3,100 F; four cycles totaling 28 minutes) In Phase I of the project, C-CAT successfully demonstrated large-scale manufacturability with a 40-inch cylinder representing the end of a nozzle extension and a 16-inch flanged cylinder representing the attach flange of a nozzle extension. These demonstrators were manufactured without spalling or delaminations. In Phase II, C-CAT worked with engine designers to develop a nozzle extension stub skirt interfaced with an Aerojet Rocketdyne RL10 engine. All objectives for Phase II were successfully met. Additional nonengine applications for the coating include thermal protection systems (TPS) for next-generation spacecraft and hypersonic aircraft.

Thompson, James↗

Liquid fluorine/hydrazine rhenium thruster update

The status of a fluorine/hydrazine thruster development program is discussed. A solid rhenium metal sea-level thrust chamber was successfully fabricated and tested for a total run duration of 1075 s with 17 starts. Rhenium fabrication methods are discussed. A test program was conducted to evaluate performance and chamber cooling. Acceptable performance was reached and cooling was adequate. A flight-type injector was fabricated that achieved an average extrapolated performance value of 3608 N-s/kg (368 lbf-s/lbm). Altitude thrust chambers were fabricated. One chamber incorporates a rhenium combustor and nozzle with an area ratio of 15:1, and a columbium nozzle extension with area ratios from 15:1 to 60:1. The other chamber was fabricated completely with a carbon/carbon composite. Because of the attributes of rhenium for use in high-temperature applications, a program to provide the materials and processes technology needed to reliably fabricate and/or repair vapor-deposited rhenium parts of relatively large size and complex shape is recommended.

Appel, M. A.↗

Molding apparatus

Apparatus for compression molding of thermosetting plastics compositions including interfitting hollow male and female components is reported. The components are adapted to be compressed to form a rocket nozzle in a cavity. A thermal jacket is provided exteriorly adjacent to the female component for circulating a thermal transfer fluid to effect curing of a thermosetting plastics material being molded. Each of the male and female components is provided with suitable inlets and outlets for circulating a thermal transfer fluid.

Heier, W. C.↗

Experimental investigation of reactive turbulent recirculating jet mixing in a dump

A ducted, subsonic, hydrogen air turbulent jet mixing flowfield was investigated, both with and without combustion. The geometric configuration was representative of a sudden expansion or dump combustor, with a central air jet surrounded by a low velocity hydrogen stream at an overall equivalence ratio of 0.12. The ratio of the duct to inner nozzle diameter was 2.5. Radial distributions of mean axial and radial velocity, axial and radial turbulent intensity, velocity cross correlation, gas composition, static temperature and total pressure, as well as axial distributions of wall static pressure, were obtained for axial stations from zero to six duct diameters from the combustor entrance. The data indicate that mixing is slower in the chemically reactive flow field than in the nonreactive flow field, and that the presence of combustion has a significant effect on the size and location of the recirculation zone within the mixing duct.

Smith, G. D.↗

Counterflow diffusion flames of hydrogen, and hydrogen plus methane, ethylene, propane, and silane vs. air - Strain rates at extinction

Five coaxial tubular opposed jet burners (OJBs) with tube diameter D(T) of 1.8-10 mm and 5 mm conical nozzles were used to form dish-shaped counterflow diffusion flames centered by opposing laminar jets of nitrogen and hydrocarbon-diluted H2 versus air in an argon-purged chamber at 1 atm. Area-averaged air jet velocities at blowoff of the central flame, U(air), characterized extinction of the airside flame as functions of input H2 concentration on the fuelside. A master plot of extensive U(air) data at blowoff versus D(T) shows that U(air) varies linearly with D(T). This and other data sets are used to find that nozzle OJB results for U(air)/diameter average 4.24 + or - 0.28 times larger than tubular OJB results for the same fuel compositions. Critical radial velocity gradients consistent with one-dimensional stagnation point boundary theory and with plug flow inputs are estimated. The results compare favorably with published numerical results based only on potential flow.

Pellett, G. L.↗