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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Methods for Modeling Particle Impingement in Solid Rocket Motor Nozzles

This paper describes several approaches available to analysts to capture the mechanical erosion effects of alumina droplet impingement in solid rocket motor nozzles. Supersonic alumina impingement is of great interest to solid rocket motor analysts, as it can be one of the main contributors to overall nozzle erosion, particularly in nozzles with highly contoured exit cones. An overview of the basic concepts of mechanical impingement is provided, along with discussion on some of the various methods available, ranging from the purely empirical to more physics-based techniques, and their respective advantages and disadvantages.

William Allen Till↗
NASA NTRS1961

Effects of Ionizing Radiation on Solid Rocket Motor Components

Problems which the solid propellant rocket engineer will encounter in designing for long-term storage in a radiation environment are discussed. A summary of present knowledge of the radiation environment is given. Mechanisms of radiation degradation and its effects on tensile properties of propellant binders are discussed qualitatively. Data from a program of irradiation of several propellants is presented. Properties of two of the propellants were changed significantly by doses of the order of 4 x 10(exp 6) rads.

RADIATION↗
NASA NTRS1980

Ice nuclei measurements from solid rocket motor effluents

The ice crystal forming nuclei (IN) measured in solid rocket motor (SRM) exhaust products is discussed in relation to space shuttle exhaust. Preliminary results from laboratory investigations and flight preparations for March 1978 Titan launch are discussed. The work necessary to provide adequate measurements of IN and cloud condensation nuclei (CCN) in the stabilized ground clouds from SRM's is studied.

Hindman, E. E., II↗
NASA NTRS1986

Block 2 Solid Rocket Motor (SRM) conceptual design study, volume 1

Segmented and monolithic Solid Rocket Motor (SRM) design concepts were evaluated with emphasis on joints and seals. Particular attention was directed to eliminating deficiencies in the SRM High Performance Motor (HPM). The selected conceptual design is described and discussed.

Source record↗
NASA NTRS2017

Arc Jet Test and Analysis of Asbestos Free Solid Rocket Motor Nozzle Dome Ablative Materials

Asbestos free solid motor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

Clayton, J. Louie↗
NASA NTRS1991

Propellant development for the Advanced Solid Rocket Motor

The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

Landers, L. C.↗
NASA NTRS1992

Trend analysis for large solid rocket motors - A program level approach

This paper defines a program-level trend analysis effort that can be applied to large solid rocket motors. The effort is especially applicable to large segmented rocket motors such as the Space Transportation System's solid rocket boosters. Five types of trend analysis are discussed for different aspects of a rocket motor program, (performance, reliability, problem, supportability, and programmatic). Ideas are offered for implementing and performing a program-level trend analysis effort by giving suggestions for selecting computer capabilities, choosing parameters, selecting statistical processes, routine screening for trends, analyzing significant trends (for example, looking for related trends and determining if adverse trends are resolved), and reporting results of trend analysis. The use of program-level trending allows for the ability to easily transfer data between organizations and easily use the data to correlate trends from the various organizations to find causal relationships. Efficiency is enhanced from upper level management to the shop floor at vendor locations by using the same trend analysis software and methodology at all organizations.

Babbitt, Norman E., III↗
NASA NTRS1992

Results of an experimental investigation of the flow field in the head-end star slot section of a Solid Rocket Motor

Cold flow tests of a one-tenth scale model based on the geometry of the Space Shuttle Solid Rocket Motor, SRM, head-end section and its single-port igniter, along with two models of a four port igniter, have been conducted. The tests were done to establish quantitative as well as qualitative data on the behavior of the flow inside the star slot during the early part of the ignition transient. This paper presents the data obtained from these tests and discusses the implications of the data with respect to ignition transients of solid rocket motors which have head-end star grains.

Ciucci, A.↗
NASA NTRS2015

Solid Rocket Motor Combustion Instability Modeling in COMSOL Multiphysics

Combustion instability modeling of Solid Rocket Motors (SRM) remains a topic of active research. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process, acoustics, and steady-state gas dynamics. Recent advances in defining the energy transport of disturbances within steady flow-fields have been applied by combustion stability modelers to improve the analysis framework [1, 2, 3]. Employing this more accurate global energy balance requires a higher fidelity model of the SRM flow-field and acoustic mode shapes. The current industry standard analysis tool utilizes a one dimensional analysis of the time dependent fluid dynamics along with a quasi-three dimensional propellant grain regression model to determine the SRM ballistics. The code then couples with another application that calculates the eigenvalues of the one dimensional homogenous wave equation. The mean flow parameters and acoustic normal modes are coupled to evaluate the stability theory developed and popularized by Culick [4, 5]. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The current study employs the COMSOL multiphysics finite element framework to model the steady flow-field parameters and acoustic normal modes of a generic SRM. The study requires one way coupling of the CFD High Mach Number Flow (HMNF) and mathematics module. The HMNF module evaluates the gas flow inside of a SRM using St. Robert's law to model the solid propellant burn rate, no slip boundary conditions, and the hybrid outflow condition. Results from the HMNF model are verified by comparing the pertinent ballistics parameters with the industry standard code outputs (i.e. pressure drop, thrust, ect.). These results are then used by the coefficient form of the mathematics module to determine the complex eigenvalues of the Acoustic Velocity Potential Equation (AVPE). The mathematics model is truncated at the nozzle sonic line, where a zero flux boundary condition is self-satisfying. The remaining boundaries are modeled with a zero flux boundary condition, assuming zero acoustic absorption on all surfaces. The results of the steady-state CFD and AVPE analyses are used to calculate the linear acoustic growth rate as is defined by Flandro and Jacob [2, 3]. In order to verify the process implemented within COMSOL we first employ the Culick theory and compare the results with the industry standard. After the process is verified, the Flandro/Jacob energy balance theory is employed and results displayed.

Fischbach, Sean R.↗
NASA NTRS2015

Solid Rocket Motor Combustion Instability Modeling in COMSOL Multiphysics

Combustion instability modeling of Solid Rocket Motors (SRM) remains a topic of active research. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process, acoustics, and steady-state gas dynamics. Recent advances in defining the energy transport of disturbances within steady flow-fields have been applied by combustion stability modelers to improve the analysis framework. Employing this more accurate global energy balance requires a higher fidelity model of the SRM flow-field and acoustic mode shapes. The current industry standard analysis tool utilizes a one dimensional analysis of the time dependent fluid dynamics along with a quasi-three dimensional propellant grain regression model to determine the SRM ballistics. The code then couples with another application that calculates the eigenvalues of the one dimensional homogenous wave equation. The mean flow parameters and acoustic normal modes are coupled to evaluate the stability theory developed and popularized by Culick. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The current study employs the COMSOL Multiphysics finite element framework to model the steady flow-field parameters and acoustic normal modes of a generic SRM. This work builds upon previous efforts to verify the use of the acoustic velocity potential equation (AVPE) laid out by Campos. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, del squared psi - (lambda/c) squared psi - M x [M x del((del)(psi))] - 2((lambda)(M)/c + M x del(M) x (del)(psi) - 2(lambda)(psi)[M x del(1/c)] = 0. with M as the Mach vector, c as the speed of sound, and λ as the complex eigenvalue. The study requires one way coupling of the CFD High Mach Number Flow (HMNF) and mathematics module. The HMNF module evaluates the gas flow inside of a SRM using St. Robert's law to model the solid propellant burn rate, slip boundary conditions, and the supersonic outflow condition. Results from the HMNF model are verified by comparing the pertinent ballistics parameters with the industry standard code outputs (i.e. pressure drop, axial velocity, exit velocity). These results are then used by the coefficient form of the mathematics module to determine the complex eigenvalues of the AVPE. The mathematics model is truncated at the nozzle sonic line, where a zero flux boundary condition is self-satisfying. The remaining boundaries are modeled with a zero flux boundary condition, assuming zero acoustic absorption on all surfaces. The one way coupled analysis is perform four times utilizing geometries determined through traditional SRM modeling procedures. The results of the steady-state CFD and AVPE analyses are used to calculate the linear acoustic growth rate as is defined by Flandro and Jacob. In order to verify the process implemented within COMSOL we first employ the Culick theory and compare the results with the industry standard. After the process is verified, the Flandro/Jacob energy balance theory is employed and results displayed.

Fischbach, S. R.↗
NASA NTRS1989

O-ring sealing verification for the space shuttle redesign solid rocket motor

As a part of the redesign of the Space Shuttle Solid Rocket Motor, the field and nozzle-to-case joints were redesigned to minimize the dynamic flexure caused by internal motor pressurization during ignition. The O-ring seals and glands for the joints were designed to accommodate both structural deflections and to promote pressure assistance. A test program was conducted to determine if a fluorocarbon elastomeric O-ring could meet this criteria in the redesigned gland. Resiliency tests were used to investigate the O-ring response to gap motion while static seal tests were used to verify design criteria of pressure assistance for sealing. All tests were conducted in face seal fixtures mounted in servo-hydraulic test machines. The resiliency of the O-ring was found to be extremely sensitive to the effects of temperature. The External Tank/Solid Rocket Booster attach strut loads had a negligible affect on the ability of the O-ring to track the simulated SRB field joint deflection. In the static pressure-assisted seal tests, as long as physical contact was maintained between the O-ring and the gland sealing surface, pressure assistance induced instantaneous sealing.

Lach, Cynthia L.↗
NASA NTRS1990

Space shuttle redesigned solid rocket motor Certificate of Qualification (COQ) data report

The Space Shuttle Redesigned Solid Rocket Motor (RSRM) Certification Program provides confidence that the RSRM and its components/subsystems meet or exceed Mission Oriented Requirements when manufactured per design requirements and specified/approved processes. Certification is based on documented results of tests, analyses, inspections, similarity, and demonstrations. Evidencing information is provided to certify that each RSRM component/subsystem satisfies design, mission related requirements and objectives.

Duersch, Fred, Jr.↗
NASA NTRS2015

An Internal Thermal Environment Model of an Aluminized Solid Rocket Motor with Experimental Validation

Due to the severity of the internal solid rocket motor (SRM) environment, very few direct measurements of that environment exist; therefore, the appearance of such data provides a unique opportunity to assess current thermal/fluid modeling capabilities. As part of a previous study of SRM internal insulation performance, the internal thermal environment of a laboratory-scale SRM featuring aluminized propellant was characterized with two types of custom heat-flux calorimeters: one that measured the total heat flux to a graphite slab within the SRM chamber and another that measured the thermal radiation flux. Therefore, in the current study, a thermal/fluid model of this lab-scale SRM was constructed using ANSYS Fluent to predict not only the flow field structure within the SRM and the convective heat transfer to the interior walls, but also the resulting dispersion of alumina droplets and the radiative heat transfer to the interior walls. The dispersion of alumina droplets within the SRM chamber was determined by employing the Lagrangian discrete phase model that was fully coupled to the Eulerian gas-phase flow. The P1-approximation was engaged to model the radiative heat transfer through the SRM chamber where the radiative contributions of the gas phase were ignored and the aggregate radiative properties of the alumina dispersion were computed from the radiative properties of its individual constituent droplets, which were sourced from literature. The convective and radiative heat fluxes computed from the thermal/fluid model were then compared with those measured in the lab-scale SRM test firings and the modeling approach evaluated.

Martin, Heath T.↗
NASA NTRS2010

Comparisons Between Stability Prediction and Measurements for the Reusable Solid Rocket Motor

The Space Transportation System has used the solid rocket boosters for lift-off and ascent propulsion over the history of the program. Part of the structural loads assessment of the assembled vehicle is the contribution due to solid rocket booster thrust oscillations. These thrust oscillations are a consequence of internal motor pressure oscillations active during operation. Understanding of these pressure oscillations is key to predicting the subsequent thrust oscillations and vehicle loading. The pressure oscillation characteristics of the Reusable Solid Rocket Motor (RSRM) design are reviewed in this work. Dynamic pressure data from the static test and flight history are shown, with emphasis on amplitude, frequency, and timing of the oscillations. Physical mechanisms that cause these oscillations are described by comparing data observations to predictions made by the Solid Stability Prediction (SSP) code.

Fischbach, Sean R.↗
NASA NTRS1995

Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

Solid State Laser Initiated Ordnance (LIO) offers new technology having potential for enhanced safety, reduced costs, and improved operational efficiency. Concerns over the absence of programmatic applications of the technology, which has prevented acceptance by flight programs, should be abated since LIO has now been operationally implemented by the Laser Initiated Ordnance Sounding Rocket Demonstration (LOSRD) Program. The first launch of solid state laser diode LIO at the NASA Wallops Flight Facility (WFF) occurred on March 15, 1995 with all mission objectives accomplished. This project, Phase 3 of a series of three NASA Headquarters LIO demonstration initiatives, accomplished its objective by the flight of a dedicated, all-LIO sounding rocket mission using a two-stage Nike-Orion launch vehicle. LIO flight hardware, made by The Ensign-Bickford Company under NASA's first Cooperative Agreement with Profit Making Organizations, safely initiated three demanding pyrotechnic sequence events, namely, solid rocket motor ignition from the ground and in flight, and flight termination, i.e., as a Flight Termination System (FTS). A flight LIO system was designed, built, tested, and flown to support the objectives of quickly and inexpensively putting LIO through ground and flight operational paces. The hardware was fully qualified for this mission, including component testing as well as a full-scale system test. The launch accomplished all mission objectives in less than 11 months from proposal receipt. This paper concentrates on accomplishments of the ordnance aspects of the program and on the program's implementation and results. While this program does not generically qualify LIO for all applications, it demonstrated the safety, technical, and operational feasibility of those two most demanding applications, using an all solid state safe and arm system in critical flight applications.

Schulze, Norman R.↗
NASA NTRS1991

The space shuttle advanced solid rocket motor: Quality control and testing

The Congressional committees that authorize the activities of NASA requested that the National Research Council (NRC) review the testing and quality assurance programs for the Advanced Solid Rocket Motor (ASRM) program. The proposed ASRM design incorporates numerous features that are significant departures from the Redesigned Solid Rocket Motor (RSRM). The NRC review concentrated mainly on these features. Primary among these are the steel case material, welding rather than pinning of case factory joints, a bolted field joint designed to close upon firing the rocket, continuous mixing and casting of the solid propellant in place of the current batch processes, use of asbestos-free insulation, and a lightweight nozzle. The committee's assessment of these and other features of the ASRM are presented in terms of their potential impact on flight safety.

Source record↗