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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1970

The aerodynamic characteristics of large angled cones with retrorockets

Analytical and experimental phases of the subject investigation are described. The analytical program for the single jet determines the terminal shock location, the jet boundary, the interface profile, the bow shock profile, the shear layer growth and the dead air region pressure. The experimental program described was conducted over the range from free stream Mach 0.4 to 2.0 at angles-of-attack up to 18 deg and at thrusting coefficients up to C sub T = T/q sub infinity A sub m = 30. Variables investigated included aeroshell angle, number of nozzles, engine thrust, size of nozzles, nozzle throttling and gas composition. The influence of these variables on the aeroshell stability, drag, and loads was determined by integrating pressure measurements on the aeroshell. The total system forces consist of components due to pure thrust and components due to pressure on the aeroshell arising from the jet-free stream interaction. Shadowgraphs provided flow field geometries which proved to be within 10% of those predicted analytically.

Jarvinen, P. O.↗
NASA NTRS1973

A high-pressure carbon dioxide gasdynamic laser

A carbon dioxide gasdynamic laser was operated over a range of reservoir pressure and temperature, test-gas mixture, and nozzle geometry. A significant result is the dominant influence of nozzle geometry on laser power at high pressure. High reservoir pressure can be effectively utilized to increase laser power if nozzle geometry is chosen to efficiently freeze the test gas. Maximum power density increased from 3.3 W/cu cm of optical cavity volume for an inefficient nozzle to 83.4 W/cu cm at 115 atm for a more efficient nozzle. Variation in the composition of the test gas also caused large changes in laser power output. Most notable is the influence of the catalyst (helium or water vapor) that was used to depopulate the lower vibrational state of the carbon dioxide. Water caused an extreme deterioration of laser power at high pressure (100 atm), whereas, at low pressure the laser for the two catalysts approached similar values. It appears that at high pressure the depopulation of the upper laser level of the carbon dioxide by the water predominates over the lower state depopulation, thus destroying the inversion.

Kuehn, D. M.↗
NASA NTRS1975

NO sub x deposited in the stratosphere by the space shuttle, phase 1

The results of calculations to determine the amount of NOx deposited in the stratosphere by space shuttle solid rocket motors (SRM) are presented. Flow properties and chemical composition in the exhaust nozzle and plume were analyzed. The nozzle calculations show that about 4.5 lbm/sec of NOx leaves the two SRMs. The total amount of NOx deposited in the stratosphere is related to the amount leaving the nozzle via an overall plume enhancement factor (OPEF), which depends upon the influence of afterburning and shocks in enhancing the exit plane NOx mole fraction. Calculations show that the OPEF is approximately 2, indicating the mass flow of NOx in the plume to be approximately l0 lbm/sec at 30 km altitude with a possible error factor of 4. For a vehicle velocity of 3750 ft/sec, therefore, the NOx deposition rate in the stratosphere is about 2.7 x 10(-3) lbm/ft.

Pergament, H. S.↗
NASA NTRS1995

1995 JANNAF Propulsion Meeting

This volume is a collection of 36 unclassified/unlimited distribution papers which were presented at the 1995 Joint Army-Navy-NASA-Air Force (JANNAF) propulsion meeting. Specific subjects discussed include the integrated High Payoff Rocket Propulsion Technology initiative, hybrid propulsion, electric propulsion, the Minuteman 2/3 missile system, slag, aluminum in propellant compositions, electric propulsion, rocket nozzle design, and tactical missiles.

Eggleston, Debra S.↗
NASA NTRS2015

Oxide_Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project

LibertyWorks®, a subsidiary of Rolls-Royce Corporation, first studied CMC (ceramic matrix composite) exhaust mixers for potential weight benefits in 2008. Oxide CMC potentially offered weight reduction, higher temperature capability, and the ability to fabricate complex-shapes for increased mixing and noise suppression. In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project (within the Integrated Systems Research Program). ERA subtasks, including those focused on CMC components, were being formulated with the goal of maturing technology from Proof of Concept Validation (Technology Readiness Level 3 (TRL 3)) to System/Subsystem or Prototype Demonstration in a Relevant Environment (TRL 6). LibertyWorks®, a subsidiary of Rolls-Royce Corporation, first studied CMC (ceramic matrix composite) exhaust mixers for potential weight benefits in 2008. Oxide CMC potentially offered weight reduction, higher temperature capability, and the ability to fabricate complex-shapes for increased mixing and noise suppression. In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project (within the Integrated Systems Research Program). ERA subtasks, including those focused on CMC components, were being formulated with the goal of maturing technology from Proof of Concept Validation (Technology Readiness Level 3 (TRL 3)) to System/Subsystem or Prototype Demonstration in a Relevant Environment (TRL 6). Oxide CMC component at both room and elevated temperatures. A TRL≈5 (Component Validation in a Relevant Environment) was attained and the CMC mixer was cleared for ground testing on a Rolls-Royce AE3007 engine for performance evaluation to achieve TRL 6.

propulsion↗
NASA NTRS2024

Hybrid Thermally Efficient Core (HyTEC) HyTEC Phase 1 – Advanced Aerodynamics Final Report

The objective of the Hybrid Thermally Efficient Core (HyTEC) – Advanced High Pressure Turbine (HPT) Aerodynamics project is to develop technology for a compact core that contributes to significant fuel burn reductions of 5-10% over current generation technologies. To accomplish this, the HPT is incorporating a range of aerodynamic features and technologies to improve component efficiency and provide favorable systems level trades. In particular, this project explored low solidity airfoils, advanced tip treatments, platform contouring, and advanced ceramic matrix composite (CMC) Stage 2 Nozzle (S2N) airfoils to eliminate the need for post throat cooling in a compact core environment. The maturation of these technology areas is expected to provide a significant improvement in component efficiencies, and consequently reductions in fuel burn, over the current state of the art (SoA). To mature these technologies to Technology Readiness Level (TRL) 4, a test campaign was performed that consisted of four tests at three facilities. TRL 3 testing was performed in the CW22 linear cascade at NASA Glenn for blade and nozzle technologies, TRL 4 nozzle testing was performed at GE Aerospace (GEA) Test Cell A8, and TRL 4 blade testing was performed at the Notre Dame Turbomachinery Laboratory (NDTL) using the Transonic Research Turbine (TRT) rig. Low solidity was successfully demonstrated to a TRL 4 level. Low solidity nozzles showed benefits in line with pre-project expectations, while low solidity blades were shown to have an aerodynamic penalty. Crucially, this program only considered the aerodynamic losses, and systems trades such as reductions in cooling flows are expected to continue to make low solidity blades a net positive. By successfully quantifying the aerodynamic performance in this project, these trades can be conducted to determine where in the engine architecture low solidity blades will contribute positively to system operation. Platform contouring was demonstrated to a TRL 4 level, with performance in line with the lower end of the pre-project expected range. Advanced tip treatments performance levels were indeterminate, showing the expected improvements to flow physics but with a performance level confounded by several rig issues including whirl mode induced variation in tip clearance. The elimination of post-throat cooling on the S2N was successfully demonstrated to provide a performance benefit, however that benefit was approximately half the level that was expected in pre-project predictions. Overall, the technology maturation plan for HyTEC Phase 1 was successful, bringing the suite of technologies to TRL 4.

High pressure turbine↗
NASA NTRS1975

Nozzle exit exhaust products from space shuttle boost vehicle (November 1973 design)

Principal exhaust species emitted at various altitudes for two trajectories of the space shuttle vehicle are presented. The exhaust composition is given for the nozzle exit plane on the basis of equilibrium chemistry. Afterburning of excess H, H2, and CO in the plume is accounted for. Species considered include HCl and Al2O3, which have been recognized as environmentally significant, as well as others such as H2O (produced by both the solid rocket motor and the orbiter main engine) which, although innocuous, may participate in subsequent chemical reactions in the atmosphere.

Source record↗
NASA NTRS1988

Standardization of the carbon-phenolic materials and processes. Vol. 1: Experimental studies

Carbon-phenolic composite materials are used as ablative material in the solid rocket motor nozzle of the Space Shuttle. The nozzle is lined with carbon cloth-phenolic resin composites. The nominal effects of the completely consumed solid propellant on the carbon-phenolic material are given. The extreme heat and erosion of the burning propellant are controlled by the carbon-phenolic composite by ablation, the heat and mass transfer process in which a large amount of heat is absorbed by sacrificially removing material from the nozzle surface. Phenolic materials ablate with the initial formation of a char. The depth of the char is a function of the heat conduction coefficient of the composite. The char layer is a very poor heat conductor so it protects the underlying phenolic composite from the high heat of the burning propellant. The nozzle component ablative liners (carbon cloth-phenolic composites) are tape wrapped, hydroclave and/or autoclave cured, machined, and assembled. The tape consists of a prepreg broadcloth. The materials flow sheet for the nozzle ablative liners is shown. The prepreg is a three component system: phenolic resin, carbon cloth, and carbon filler. This is Volume 1 of two, Experimental Studies.

Hall, William B.↗
NASA NTRS1989

High temperature insulation barrier composite

A composite material suitable for providing insulation for the nozzle structure of the Space Shuttle and other similar surfaces is disclosed. The composite layer is comprised of an outer skin layer of nickel chromium and an interleaved inner region comprising a top layer of nickel chromium foil which acts as a primary convective shield. There are at least two layers of alumina batting adjacent to the layers of silicon carbide fabric. An additional layer of nickel chromium foil is used as a secondary convective shield. The composite is particularly advantageous for use as nozzle insulation because of its ability to withstand high reentry temperatures, its flexibility, oxidation resistance, low conductivity, and light weight.

Onstott, Joseph W.↗
NASA NTRS1995

Mechanical Properties of Continuous Fiber Reinforced Zirconium Diboride Matrix Composites

Continuous fiber reinforced zirconium diboride matrix composites, SCS-9a-(RBSiCZrB2)matrix, are being developed for leading edge, rocket nozzle and turbine engine applications. Recently, the composite materials have been characterized for tensile properties to 1250 C, the highest temperature tested. The tensile properties are fiber dominated as the matrix is microcracked on fabrication, but favorable failure characteristic are observed. Compression and shear mechanical testing results will be reported if completed. The effects of fiber volume fraction and matrix density on mechanical properties will be discussed. The target applications of the materials will be discussed. Specific testing being performed towards qualification for these applications will be included.

Stuffle, Kevin↗
NASA NTRS2007

Elevated Temperature Fatigue Endurance of Three Ceramic Matrix Composites

High-cycle fatigue endurance of three candidate materials for the acoustic liners of the Enabling Propulsion Materials Nozzle Program was investigated. The ceramic matrix composite materials investigated were N720/AS (Nextel 720, 3M Corporation), Sylramic S200 (Dow Corning), and UT 22. High-cycle fatigue tests were conducted in air at 910 C on as-machined specimens and on specimens subjected to tensile cyclic load excursions every 160 hr followed by thermal exposure at 910 C in a furnace up to total exposure times of 2066 and 4000 hr. All the fatigue tests were conducted in air at 100 Hz with a servohydraulic test machine. In the as-machined condition, among the three materials investigated only the Sylramic S200 exhibited a deterministic type of high-cycle fatigue behavior. Both the N720/AS and UT-22 exhibited significant scatter in the experimentally observed high-cycle fatigue lives. Among the thermally exposed specimens, N720/AS and Sylramic S200 materials exhibited a reduction in the high-cycle fatigue lives, particularly at the exposure time of 4000 hr.

Kalluri, Sreeramesh↗
NASA NTRS2015

A Fully Non-Metallic Gas Turbine Engine Enabled by Additive Manufacturing

In a NASA Aeronautics Research Institute (NARI) sponsored program entitled "A Fully Non-Metallic Gas Turbine Engine Enabled by Additive Manufacturing," evaluation of emerging materials and additive manufacturing technologies was carried out. These technologies may enable fully non-metallic gas turbine engines in the future. This paper highlights the results of engine system trade studies which were carried out to estimate reduction in engine emissions and fuel burn enabled due to advanced materials and manufacturing processes. A number of key engine components were identified in which advanced materials and additive manufacturing processes would provide the most significant benefits to engine operation. In addition, feasibility of using additive manufacturing technologies to fabricate gas turbine engine components from polymer and ceramic matrix composite were demonstrated. A wide variety of prototype components (inlet guide vanes (IGV), acoustic liners, engine access door, were additively manufactured using high temperature polymer materials. Ceramic matrix composite components included first stage nozzle segments and high pressure turbine nozzle segments for a cooled doublet vane. In addition, IGVs and acoustic liners were tested in simulated engine conditions in test rigs. The test results are reported and discussed in detail.

Ceramic Matrix Composites↗
NASA NTRS2015

A Fully Non-Metallic Gas Turbine Engine Enabled by Additive Manufacturing

In a NASA Aeronautics Research Institute (NARI) sponsored program entitled "A Fully Non-Metallic Gas Turbine Engine Enabled by Additive Manufacturing", evaluation of emerging materials and additive manufacturing technologies was carried out. These technologies may enable fully non-metallic gas turbine engines in the future. This paper highlights the results of engine system trade studies which were carried out to estimate reduction in engine emissions and fuel burn enabled due to advanced materials and manufacturing processes. A number of key engine components were identified in which advanced materials and additive manufacturing processes would provide the most significant benefits to engine operation. In addition, feasibility of using additive manufacturing technologies to fabricate gas turbine engine components from polymer and ceramic matrix composite were demonstrated. A wide variety of prototype components (inlet guide vanes (IGV), acoustic liners, engine access door) were additively manufactured using high temperature polymer materials. Ceramic matrix composite components included first stage nozzle segments and high pressure turbine nozzle segments for a cooled doublet vane. In addition, IGVs and acoustic liners were tested in simulated engine conditions in test rigs. The test results are reported and discussed in detail.

Additive Manufacturing↗
NASA NTRS1999

Aerothermo-Structural Analysis of Low Cost Composite Nozzle/Inlet Components

This research is a cooperative effort among the Turbomachinery and Propulsion Division of NASA Glenn, CCMR of NC A&T State University, and the Tuskegee University. The NC A&T is the lead center and Tuskegee University is the participating institution. Objectives of the research were to develop an integrated aerodynamic, thermal and structural analysis code for design of aircraft engine components, such as, nozzles and inlets made of textile composites; conduct design studies on typical inlets for hypersonic transportation vehicles and setup standards test examples and finally manufacture a scaled down composite inlet. These objectives are accomplished through the following seven tasks: (1) identify the relevant public domain codes for all three types of analysis; (2) evaluate the codes for the accuracy of results and computational efficiency; (3) develop aero-thermal and thermal structural mapping algorithms; (4) integrate all the codes into one single code; (5) write a graphical user interface to improve the user friendliness of the code; (6) conduct test studies for rocket based combined-cycle engine inlet; and finally (7) fabricate a demonstration inlet model using textile preform composites. Tasks one, two and six are being pursued. Selected and evaluated NPARC for flow field analysis, CSTEM for in-depth thermal analysis of inlets and nozzles and FRAC3D for stress analysis. These codes have been independently verified for accuracy and performance. In addition, graphical user interface based on micromechanics analysis for laminated as well as textile composites was developed. Demonstration of this code will be made at the conference. A rocket based combined cycle engine was selected for test studies. Flow field analysis of various inlet geometries were studied. Integration of codes is being continued. The codes developed are being applied to a candidate example of trailblazer engine proposed for space transportation. A successful development of the code will provide a simpler, faster and user-friendly tool for conducting design studies of aircraft and spacecraft engines, applicable in high speed civil transport and space missions.

Shivakumar, Kuwigai↗
NASA NTRS1972

Molecular-beam gas-sampling system

A molecular beam mass spectrometer system for rocket motor combustion chamber sampling is described. The history of the sampling system is reviewed. The problems associated with rocket motor combustion chamber sampling are reported. Several design equations are presented. The results of the experiments include the effects of cooling water flow rates, the optimum separation gap between the end plate and sampling nozzle, and preliminary data on compositions in a rocket motor combustion chamber.

Young, W. S.↗
NASA NTRSDate not supplied

3D Scanning Laboratory

The 3D Scanning Lab uses a number of technologies to capture 3D surface data. Those technologies are structured light scanning, discrete point measurement photogrammetry and Light Detecting and Ranging(LIDAR). Structured light is a non-contact optical technique used to capture as built surface geometries quickly and accurately. LIDAR is a laser scanning technique used to capture 3D surfaces. And photogrammetry uses a certified DSLR camera to measure discrete registered and non-registered points in 3D space. MSFC’s 3D Scanning Team uses 3D scanning to provide hardware inspections, reverse engineered CAD models, manufacturing/process development and digital assemblies. The inspection process compares the captured data to CAD, providing a color plot detailing the deviations of the scan data to the CAD model. Dimensional measurements and GD&T can also be interrogated. Inspection can provide scan-to-scan data comparisons which is ideal for assessing tested hardware, e.g. comparing data captured before and after hardware testing. Reverse engineering is used to develop CAD models from scan data. This is useful for heritage hardware studies, developing building or structural models and inputs for simulation studies and design efforts. The lab uses the 3D surface scan data to drive manufacturing/machining processes, termed match machining. This technique has been used to modify nozzles, injectors, additive manufactured parts, composite barrel sections, etc. Parts of an assembly can be scanned and assembled in a virtual environment providing an accurate 3D representation of the assembled hardware in its as-scanned state. This can provide detailed information of internal components that would otherwise not be accessible.

Advanced Manufacturing↗
NASA NTRS1999

NASA's Advanced Space Transportation Program: A Materials Overview

The realization of low-cost assess to space is one of NASA's three principal goals or "pillars" under the Office of Aero-Space Technology. In accordance with the goals of this pillar, NASA's primary space transportation technology role is to develop and demonstrate next-generation technologies to enable the commercial launch industry to develop full-scale, low cost, highly reliable space launchers. The approach involves both ground-based technology demonstrations and flight demonstrators, including the X-33, X-34, Bantam, Reusable Launch Vehicle (RLV), and future experimental vehicles. Next generation space transportation vehicles and propulsion systems will require the development and implementation of advanced materials and processes. This presentation will provide an overview of advanced materials efforts which are focused on the needs of next generation space transportation systems. Applications described will include ceramic matrix composite (CMC) integrally bladed turbine disk (blisk); actively cooled CMC nozzle ramp for the aerospike engine; ablative thrust chamber/nozzle; and metal matrix composite turbomachinery housings.

Clinton, R. G., Jr.↗