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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Analyses and Methods of Solid Rocket Motor Material Irradiation at Marshall Space Flight Center

The search for life on other worlds is among humanity’s greatest endeavors. Europa represents the most probable location to discover extraterrestrial life in our solar system, owing to its surface composition of ice covering a liquid water ocean, warmed by the tidal forces of its orbit around Jupiter. Unfortunately, the Jovian system hosts the most intense planetary radiation environment in the solar system due to the charged particles, namely electrons and protons, trapped by Jupiter’s immense magnetic field. Any mission that attempts to approach or land on Europa must survive this radiation environment [Hand et al, 2017]. Radiation effects were identified as a priority risk to the successful development of a de-orbit stage and solid rocket motor (SRM) early in the Europa Lander De-orbit Stage project concept. The effects of primary concern tend to occur very near the outer surface of the SRM. The charged particles deposit their energy quickly and are mostly stopped in the outer metallic case, but a significant portion of radiation penetrates through the bondline and outer propellant regions. High doses of ionizing radiation are known to cause significant changes to mechanical properties of many materials, especially polymers. For polymers such as the rubber-like materials (elastomers) in a solid rocket motor, the primary damage mechanism is known as cross-linking, in which ionization causes the restructuring of the matrix of long polymer chains. Ionization energy breaks the long polymer chains and allows formation of new cross-linked bonding sites. This hardens and often strengthens the polymer, but at the cost of decreased flexibility (or modulus). Propellant, insulation, liner, and pyrotechnic materials were identified as higher risk items, and so were irradiated at Marshall Space Flight Center (MSFC) for investigation of changes in mechanical and ballistic properties. This process required significant levels of analysis to evaluate how the radiation environment evolves within the spacecraft during the mission, and also to evaluate how dose is delivered into test articles within the irradiating facilities.

Caffrey, Jarvis↗
NASA NTRS2020

System Engineering a Solar Thermal Propulsion Mission Concept for Rapid Interstellar Medium Access

The interstellar medium (ISM) represents the next frontier in space exploration, with many new discoveries to be made. The challenge, being so far away from Earth, the ISM requires many decades to reach. To advance our knowledge of what exists beyond our Solar System, new approaches for rapid access are required. One such approach is solar thermal propulsion (STP). The approach uses several Venus and Earth gravity assists to fly to Jupiter and use its gravity well to dive towards the Sun. Approaching within 3 solar radii a perihelion burn would be performed, maximizing the spacecraft’s ΔV to achieve high Solar System escape velocities. A unique aspect of the STP mission concept is that the Sun is not only used as a gravity well for an Oberth manoeuvre, but also to heat the fuel to ultra-high-temperatures (>3000 K), enabling a monopropellant burn with high specific impulse (Isp). Prior preliminary studies indicated escape velocities of over 20 astronomical unit (AU)/yr would be possible. An in-depth modelling exercise was undertaken to determine how such a system would perform. The model in this paper showed the current STP design is capable of providing just under 9±1 AU/yr, but there are many technology developments that could increase escape velocity. The technologies vary from items that could be implemented in the near term, like turbo-pumps driven by the hydrogen, to items requiring more extensive development programs like thin coatings which do not erode in superheated hydrogen. After reviewing the STP approach, and comparing it to a solid rocket motor (SRM), it was found that with currently available technology, SRM outperforms STP with an escape velocity of approximately 10-12 AU/yr. However, future advances in heat exchanger lining materials, turbo pumps, and advanced heat exchanger geometries may enable solar thermal propulsion to provide higher escape velocities, providing one of the fastest ways to exit the solar system. Ultimately, if all technology paths could be implemented with minimal side effects, the performance in a best-case scenario could reach up to 16 AU/yr.

Alkalai, Leon↗
NASA NTRSDate not supplied

SLS Ignition Overpressure-Sound Suppression System Performance Evaluated Against Historical Configurations

During the start-up of a number of launch vehicles that include solid rocket motor (SRM), the ignition transient and acoustic environments are mitigated by the implementation of a water spray system located immediately below the SRM nozzle exit plane (NEP). For NASA’s Space Launch System (SLS), this water system is referred to as the Ignition Overpressure / Sound Suppression (IOP/SS) system. The SLS Induced Environments (IE) technical discipline conducted a comprehensive evaluation of the design and as-tested performance of the IOP/SS water that will operate underneath both Boosters during the Artemis I launch. As part of this evaluation, flow rates and imagery from a number of integrated launch pad / mobile launcher IOP/SS flow tests were studied. Additional insight was leveraged from the Shuttle heritage IOP/SS system that includes data and imagery from a number of Flight Readiness Firings (FRF) and water flow tests. Lastly, the IE study included a qualitive comparison of the Shuttle and SLS systems to the equivalent water flow systems for Titan and Atlas V and determined that the NASA water flow systems are substantially different than those supporting other launch vehicles.

Ignition Overpressure↗
NASA NTRSDate not supplied

SLS Ignition Overpressure-Sound Suppression System Performance Evaluated Against Historical Configurations

During the start-up of a number of launch vehicles that include solid rocket motor (SRM), the ignition transient and acoustic environments are mitigated by the implementation of a water spray system located immediately below the SRM nozzle exit plane (NEP). For NASA’s Space Launch System (SLS), this water system is referred to as the Ignition Overpressure / Sound Suppression (IOP/SS) system. The SLS Induced Environments (IE) technical discipline conducted a comprehensive evaluation of the design and as-tested performance of the IOP/SS water that will operate underneath both Boosters during the Artemis I launch. As part of this evaluation, flow rates and imagery from a number of integrated launch pad / mobile launcher IOP/SS flow tests were studied. Additional insight was leveraged from the Shuttle heritage IOP/SS system that includes data and imagery from a number of Flight Readiness Firings (FRF) and water flow tests. Lastly, the IE study included a qualitive comparison of the Shuttle and SLS systems to the equivalent water flow systems for Titan and Atlas V and determined that the NASA water flow systems are substantially different than those supporting other launch vehicles.

Ignition Overpressure↗
NASA NTRS1972

Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

The design, development, production, and launch support analysis for determining the solid propellant rocket engine to be used with the space shuttle are discussed. Specific program objectives considered were: (1) definition of engine designs to satisfy the performance and configuration requirements of the various vehicle/booster concepts, (2) definition of requirements to produce booster stages at rates of 60, 40, 20, and 10 launches per year in a man-rated system, and (3) estimation of costs for the defined SRM booster stages.

Source record↗
NASA NTRS1972

Aerodynamic stability and drag characteristics of a parallel burn/SRM ascent configuration (M equals 0.6 to 4.96)

Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch trisonic wind tunnel during April 1972 on a 0.004-scale model of a solid rocket motor version of the space shuttle ascent configuration. The configuration consisted of a parallel burn solid rocket motor booster on an external HO centerline tank orbiter. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 deg to +10 deg at zero degrees sideslip and over a sideslip range from -10 deg to +10 deg at zero degrees angle of attack. Mach numbers ranged from 0.6 to 4.96. The purpose of the test was to determine the performance and stability characteristics of the complete ascent configuration and buildup, and to determine the effects of variations in HO tank and SRM nose shaping, orbiter incidence and position, and position of the solid rocket motors.

Sims, F.↗
NASA NTRS1972

Orbiter aborts from boost: Presimulation report

A description of a hybrid simulation of the 040C orbiter aborting from boost to specified landing site is provided. The simulation starts when the abort is initiated and continues until a terminal energy state (associated with the selected landing site) is reached. At abort it is assumed that all SRM's are jettisoned with the external tank remaining with the orbiter. The simulation described has six degrees of freedom with the vehicle simulated as a rigid body. A conventional form of autopilot is provided to control engine gimbaling during powered flight. An ideal form of an autopilot is provided to test conventional autopilot function and provide pseudo RCS function during coasting flight. The simulation is proposed to provide means for studies of abort guidance function and to gain information concerning ability to control the abort trajectory.

Backman, H. D.↗
NASA NTRS1972

Solid rocket motor cost model

A systematic and standardized procedure for estimating life cycle costs of solid rocket motor booster configurations. The model consists of clearly defined cost categories and appropriate cost equations in which cost is related to program and hardware parameters. Cost estimating relationships are generally based on analogous experience. In this model the experience drawn on is from estimates prepared by the study contractors. Contractors' estimates are derived by means of engineering estimates for some predetermined level of detail of the SRM hardware and program functions of the system life cycle. This method is frequently referred to as bottom-up. A parametric cost analysis is a useful technique when rapid estimates are required. This is particularly true during the planning stages of a system when hardware designs and program definition are conceptual and constantly changing as the selection process, which includes cost comparisons or trade-offs, is performed. The use of cost estimating relationships also facilitates the performance of cost sensitivity studies in which relative and comparable cost comparisons are significant.

Harney, A. G.↗
NASA NTRS1972

Development of techniques and associated instrumentation for high temperature emissivity measurements

The calibrating and testing procedures for the radiometric emittance measurement apparatus are described. Test results are given for the NBS platinum-rhodium standard (SRM No. 1409), the NBS oxidized Kanthal standard, and the oxidized Rene 41 materials. Results indicate that with the platinum-rhodium standard the measurement system compared with the NBS data to within 0.02 for both total and spectral normal emitttance. Data from the Kanthal strip was consistently higher than NBS values. For oxidized Rene 41, no significant changes in as prepared emittance was observed. The emittance was stable for temperatures up to 1150 K.

Cunnington, G. R.↗
NASA NTRS1972

Technical report analysis and design: Study of solid rocket motors for a space shuttle booster, volume 2, book 1, supplement 1

An analysis and design effort was conducted as part of the study of solid rocket motor for a space shuttle booster. The 156-inch-diameter, parallel burn solid rocket motor was selected as its baseline because it is transportable and is the most cost-effective, reliable system that has been developed and demonstrated. The basic approach was to concentrate on the selected baseline design, and to draw from the baseline sufficient data to describe the alternate approaches also studied. The following conclusions were reached with respect to technical feasibility of the use of solid rocket booster motors for the space shuttle vehicle: (1) The 156-inch, parallel-burn baseline SRM design meets NASA's study requirements while incorporating conservative safety factors. (2) The solid rocket motor booster represents a cost-effective approach. (3) Baseline costs are conservative and are based on a demonstrated design. (4) Recovery and reuse are feasible and offer substantial cost savings. (5) Abort can be accomplished successfully. (6) Ecological effects are acceptable.

Source record↗
NASA NTRS1973

Some environmental considerations relating to the interaction of the solid rocket motor exhaust with the atmosphere: Predicted chemical composition of exhaust species and predicted conditions for the formation of HCl aerosol

The exhaust products of a solid rocket motor using as propellant 14% binder, 16% aluminum, and 70% (wt) ammonium perchlorate consist of hydrogen chloride, water, alumina, and other compounds. The equilibrium and some frozen compositions of the chemical species upon interaction with the atmosphere were computed. The conditions under which hydrogen chloride interacts with the water vapor in humid air to form an aerosol containing hydrochloric acid were computed for various weight ratios of air/exhaust products. These computations were also performed for the case of a combined SRM and hydrogen-oxygen rocket engine. Regimes of temperature and relative humidity where this aerosol is expected were identified. Within these regimes, the concentration of HCL in the aerosol and weight fraction of aerosol to gas phase were plotted. Hydrochloric acid aerosol formation was found to be particularly likely in cool humid weather.

Rhein, R. A.↗
NASA NTRS1972

Space shuttle system program definition. Volume 2: Technical report

The Phase B Extension of the Space Shuttle System Program Definition study was redirected to apply primary effort to consideration of space shuttle systems utilizing either recoverable pressure fed liquids or expendable solid rocket motor boosters. Two orbiter configurations were to be considered, one with a 15x60 foot payload bay with a 65,000 lb, due East, up-payload capability and the other with a 14x45 payload bay with 45,000 lb, of due East, up-payload. Both were to use three SSME engines with 472,000 lb of vacuum thrust each. Parallel and series burn ascent modes were to be considered for the launch configurations of primary interest. A recoverable pump-fed booster is included in the study in a series burn configuration with the 15x60 orbiter. To explore the potential of the swing engine orbiter configuration in the pad abort case, it is included in the study matrix in two launch configurations, a series burn pressure fed BRB and a parallel burn SRM. The resulting matrix of configuration options is shown. The principle objectives of this study are to evaluate the cost and technical differences between the liquid and solid propellant booster systems and to assess the development and operational cost savings available with a smaller orbiter.

Source record↗