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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Manufacture and static firing of X259-E6 rocket motor serial number XJ04/0001

A single motor was cast and static fired to demonstrate the performance of high energy crosslinked double base (XLDB) propellant in standard X259 rocket motor hardware. Prior to motor fabrication, the motor was analyzed to predict the results of static firing the X259 motor loaded with XLDB propellant. As a result of the analyses, a forward dome shrinkage liner was added to the design. With this design change it was determined that adequate margins of safety existed. The motor, designated the X259-E6 model with serial number XJ04/0001, was fabricated using a slurry-casting technique and was assembled with a standard X259-B4 nozzle which had the nozzle throat machined to a smaller inside diameter than the B4 model and the exit cone cut short for Bacchus Works altitude expansion. The motor was static fired on 20 February 1974 with the nozzle failing during motor operation. Nozzle failure was attributed to spalling of the throat material leading to complete nozzle break-up. However, the propellant functioned as predicted in the motor chamber, ignition was normal, and char and erosion of the internal insulator were as expected.

Robertson, D. R.↗
NASA NTRS1988

Test stand for Titan 34D SRM static firing

An existing liquid engine test stand at the AF Astronautics Laboratory was refurbished and extensively modified to accommodate the static firing of the Titan 34D solid rocket motor (SRM) in the vertical nozzle down orientation. The main load restraint structure was designed and built to secure the SRM from lifting off during the firing. In addition, the structure provided weather protection, temperature conditioning of the SRM, and positioning of the measurement and recording equipment. The structure was also used for stacking/de-stacking of SRM segments and other technological processes. The existing stand, its foundation and anchorage were thoroughly examined and reanalyzed. Necessary stand modifications were carried out to comply with the requirements of the Titan 34D SRM static firing.

Glozman, Vladimir↗
NASA NTRS1972

The geometry and physical properties of exhaust clouds generated during the static firing of S-1C and S-2 rocket engines

An experimental program was conducted during the static firing of the S-1C stage 13, 14, and 15 rocket engines and the S-2 stage 13, 14, and 15 rocket engines. The data compiled during the experimental program consisted of photographic recordings of the time-dependent growth and diffusion of the exhaust clouds, the collection of meteorological data in the ambient atmosphere, and the acquisition of data on the physical structure of the exhaust clouds which were obtained by flying instrumented aircraft through the clouds. A new technique was developed to verify the previous measurements of evaporation and entrainment of blast deflector cooling water into the cloud. The results of the experimental program indicate that at the lower altitudes the rocket exhaust cloud or plume closely resembles a free-jet type of flow. At the upper altitudes, where the cloud is approaching an equilibrium condition, structure is very similar to a natural cumulus cloud.

Forbes, R. E.↗
NASA NTRS1988

Test stand design and analysis for Titan 34D static firing

The Test Stand 1C at the Air Force Astronautic Laboratory (AFAL), Edwards AFB, Calif., was originally designed for F-1 liquid propellant engines. The stand was modified to handle a five and one half segment Solid Rocket Motor (SRM) Titan 34D vertical static firing test. The main design drivers and analysis effort utilized in the modification of the test stand are discussed.

Pazargadi, Shayan↗
NASA NTRS1989

Space shuttle Production Verification Motor 1 (PV-1) static fire

All inspection and instrumentation data indicate that the PV-1 static test firing conducted 18 Aug. 1988 was successful. With the exception of the intentionally flawed joints and static test modifications, PV-1 was flight configuration. Fail-safe flaws guaranteeing pressure to test the sealing capability of primary O-rings were included in the aft field joint, case-to-nozzle joint, and nozzle internal Joint 5. The test was conducted at ambient conditions, with the exception of the field joints and case/nozzle joints which were maintained at a minimum of 75 F. Ballistics performance values were within specification requirements. The PV-1 motor exhibited chamber pressure oscillations similar to previously tested Space Shuttle redesigned solid rocket motors, particularly QM-7. The first longitudinal mode oscillations experienced by PV-1 were the strongest ever measured in a Space Shuttle motor. Investigation into this observation is being conducted. Joint insulation performed as designed with no evidence of gas flow within unflawed forward field joints. The intentionally flawed center and aft case field joint insulation performance was excellent. There was no evidence of hot gas past the center field joint capture feature O-ring, the case-to-nozzle joint primary O-ring, or the aft field joint primary O-ring. O-ring seals and barriers with assured pressure at the flaws showed erosion and heat effect, but all sealed against passage of hot gases with the exception of the aft field joint capture feature O-ring. There was no evidence of erosion, heat effect, or blowby on any O-ring seals or barriers at the unflawed joints. Nozzle performance was nominal with typical erosion. Post-test examination revealed that the forward nose ring was of the old high performance motor design configuration with the 150-deg ply angle. All nozzle components remained intact for post-test evaluation. The thrust vector control system operated correctly. The water deluge system, CO2 quench, and other test equipment performed as planned during all required test operations.

Source record↗
NASA NTRS1995

Plume Particle Collection and Sizing from Static Firing of Solid Rocket Motors

Thermal radiation from the plume of any solid rocket motor, containing aluminum as one of the propellant ingredients, is mainly from the microscopic, hot aluminum oxide particles in the plume. The plume radiation to the base components of the flight vehicle is primarily determined by the plume flowfield properties, the size distribution of the plume particles, and their optical properties. The optimum design of a vehicle base thermal protection system is dependent on the ability to accurately predict this intense thermal radiation using validated theoretical models. This article describes a successful effort to collect reasonably clean plume particle samples from the static firing of the flight simulation motor (FSM-4) on March 10, 1994 at the T-24 test bed at the Thiokol space operations facility as well as three 18.3% scaled MNASA motors tested at NASA/MSFC. Prior attempts to collect plume particles from the full-scale motor firings have been unsuccessful due to the extremely hostile thermal and acoustic environment in the vicinity of the motor nozzle.

Sambamurthi, Jay K.↗
NASA NTRS1995

Plume Diagnostics of the RSRM Static Firings for the Pressure Perturbation Studies

During the STS-54 launch (RSRM-29), the right hand solid rocket motor experienced a 13.9 psi chamber pressure perturbation at 67 seconds into the motor operation. This pressure augmentation equated to a thrust change of 51 klb. Concerns were raised regarding the adverse effects of this thrust imbalance on the shuttle system and the overall thrust into the external tank structural elements. Pressure perturbations have been observed in solid rocket motors due to expulsion of igniter or insulation materials; the motor thrust during such events drop abruptly before rising. However, the RSRM motors do not exhibit such behavior during the large chamber pressure perturbation events. Several scenarios were investigated to explain these pressure perturbations in the RSRM motors based on a fault tree developed after STS-54. Of these, the expulsion of the slag accumulated in the submerged nozzle region appeared to be the most plausible scenario to explain the observations. Slag is a natural combustion product of aluminized solid rocket motors. The RSRM propellant contains 16% by weight of aluminum. Any ejection of this slag mass during nozzle vectoring or other side loads on the motor will result in the chamber pressure perturbation. Two RSRM static firings were instrumented extensively to further understand the slag expulsion phenomenon in the RSRM and the associated pressure perturbations.

Mathias, Edward C.↗
NASA NTRS1995

Plume diagnostics of SRM static firings for pressure perturbation studies

During the shuttle launches, the solid rocket motors (SRM) occasionally experience pressure perturbations (8-13 psi) between 65 and 75 seconds into the motor burn time. The magnitudes of these perturbations are very small in comparison with the operating motor chamber pressure, which is over 600 psi during this time frame. These SRM pressure perturbations are believed to be caused primarily by the expulsion of slag (aluminum oxide). Two SRM static tests, TEM-11 and FSM-4, were instrumented extensive]y for the study of the phenomenon associated with pressure perturbations. The test instrumentation used included nonintrusive optical and infrared diagnostics of the plume, such as high-speed photography, radiometers, and thermal image cameras. Results from all these nonintrusive observations strongly support the scenario that the pressure perturbation event in the shuttle SRM is caused primarily by the expulsion of molten slag. The slag was also expelled preferentially near the bottom of the nozzle due to slag accumulation at the bottom of the aft end of the horizontally oriented motor.

Sambamurthi, J. K.↗
NASA NTRS1978

Quick look test report: MPT static firing no. 2 test MPT-S2

The three engine cluster was fired at 70 percent power level for a nominal 15 seconds to evaluate the integrated performance of the main propulsion system. Engine ignition occurred at approximately 1403 with the planned mainstage duration achieved for all three engines. Operation of all systems was as expected with the exception of the recirculation pumps. The pumps were started while the propellant loading was in fast fill, but they cavitated and lost head at the termination of fast fill. The pumps were subsequently restarted after pressurizing the tank and draining back propellant to get good quality. Post test inspection of the engines revealed some discoloration on the inside of the thrust chamber and distorted drain lines for engine #2.

Source record↗
NASA NTRS1980

Pogo summary report main propulsion test static firings 1-7 for shuttle development flight instrumentation

Problems concerning the shuttle main propulsion system Polar Orbit Geophysical Observatory (POGO) instrumentation and the actions taken to correct them are summarized. Investigations and analyses appear to be providing solutions to correct the majority of questionable measurements. Corrective action in the handling of cables and connectors should increase the POGO measurement quality. Unacceptable levels of very low frequency noise and data level shifts may be related to test stand grounding configuration, but further investigation is required.

Haddick, C. M., Jr.↗
NASA NTRS1995

Plume particle collection and sizing from static firing of solid rocket motors

A unique dart system has been designed and built at the NASA Marshall Space Flight Center to collect aluminum oxide plume particles from the plumes of large scale solid rocket motors, such as the space shuttle RSRM. The capability of this system to collect clean samples from both the vertically fired MNASA (18.3% scaled version of the RSRM) motors and the horizontally fired RSRM motor has been demonstrated. The particle mass averaged diameters, d43, measured from the samples for the different motors, ranged from 8 to 11 mu m and were independent of the dart collection surface and the motor burn time. The measured results agreed well with those calculated using the industry standard Hermsen's correlation within the standard deviation of the correlation . For each of the samples analyzed from both MNASA and RSRM motors, the distribution of the cumulative mass fraction of the plume oxide particles as a function of the particle diameter was best described by a monomodal log-normal distribution with a standard deviation of 0.13 - 0.15. This distribution agreed well with the theoretical prediction by Salita using the OD3P code for the RSRM motor at the nozzle exit plane.

Sambamurthi, Jay K.↗
NASA NTRS1972

Structural damage claims resulting from acoustic environments developed during static test firing of rocket engines

During static testing of multi-million pound thrust rocket engines areas adjacent to the test site have been subjected to the noise generated by rocket engines. Structural damage claims and subjective complaints were filed by those who alleged that the noise levels were excessive. The statistical analysis of these claims and complaints which were filed during these rocket engine development programs led to the determination of a relationship between claims and overall sound pressure level. Community exposure criteria are then assessed based on what can be considered allowable acoustic environments from large rocket engines.

Guest, S. H.↗
NASA NTRS1978

Electrets used to measure exhaust cloud effluents from Solid Rocket Motor (SRM) during demonstration model (DM-2) static test firing

Electrets were compared with fixed flow samplers during static test firing. The measurement of the rocket exhaust effluents by samplers and electrets indicated that the Solid Rocket Motor had no significant effect on the air quality in the area sampled. The results show that the electrets (a passive device which needs no power) can be used effectively alongside existing measuring devices (which need power). By placing electrets in areas where no power is available, measurements may be obtained. Consequently, it is a valuable complementary instrument in measuring rocket exhaust effluents in areas where other measuring devices may not be able to assess the contaminants.

Susko, M.↗