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Lockheed Martin was requested to develop a closed loop CIL system following the Challenger accident. The system that was developed has proven to be very robust with minimal problems since implementation, having zero escapes in the last 7 years (27 External Tanks). We are currently investigating expansion of the CIL Closed Loop system to include "MI" CILs.
The NASA Glenn Research Center at Lewis Field has taken the lead in the development of practical densified cryogenic propellants for launch vehicle applications. The technology of subcooling cryogenic propellants below their normal boiling point to produce a denser fluid is one of the key process technologies necessary to meet the challenge of single-stage-to-orbit and reusable launch vehicles. Densified propellants are critical to lowering launch costs because they enable more propellant to be packed into a given unit volume, thus improving the performance by reducing the overall size and weight of the launch vehicle. This two-pronged research and test program has evolved into (1) conducting tank loading tests using densified liquid hydrogen and (2) developing two large-scale propellant densification systems that will be performance tested next year at Glenn. The propellant-loading test program was undertaken at Glenn in coordination with Lockheed Martin Michoud Space Systems. In this testing, the liquid hydrogen recirculation and densification process was simulated, and the thermal stratification of the densified propellant was recorded throughout the tank. The test article was a flight-weight tank constructed from composite materials similar to those to be used on the X-33 launch vehicle. The tank geometry as designed by Lockheed Martin had two cylindrical lobes with a center septum. Liquid hydrogen flow rate, pressure data, and temperature data plotted over time were collected while the subscale tank was filled with 27 R (15 K) densified liquid hydrogen propellant. This testing has validated mathematical models and demonstrated the readiness of densified propellant technology for near-term use. It marks the first time that such a process has been carried out with a multiple-lobe, flight-similar tank. Glenn researchers have also been working on providing a process and critical test data for the continuous production of densified liquid hydrogen (LH2) and densified liquid oxygen (LO2). Each densification production process uses a high-efficiency, subatmospheric boiling bath heat exchanger to cool the working fluid. A near triple-point hydrogen boiling bath is used to condition and subcool hydrogen to 27 R (15 K), and a nitrogen boiling bath is used to cool the liquid oxygen to 120 R (66.7 K). Multistage centrifugal compressors operating at cryogenic inlet conditions maintain the heat exchanger bath vapor pressure below 1 atm. The LO2 propellant densification unit shown in the photograph has a 30 lb/sec capacity, whereas the LH2 unit was designed to process 8 lb/sec of propellant. Each densification unit will be transported to Glenn's South Forty area after all fabrication work is completed sometime late next year. There the LO2 and LH2 densifier performance tests will be conducted with another larger Lockheed Martin tank designated the Structural Test Article (STA). This liquid oxygen tank is a full-scale, flight-weight, prototype aluminum tank designed for the X-33. It has a capacity of 20,000 gallons of LO2. The tank loading and recirculation testing planned for next year with STA will provide the data necessary for full-scale development of propellant densification technology.
During the quarter hour of its operational life, the External Tank must provide the propellants and propellant conditioning for the Orbiter Main Engines, provide the structural support for the two Solid Rocket Boosters, withstand the natural and induced environments, and effect a safe disposal. The External Tank is twice the size of the S-II Stage used in the Saturn V Booster for the Apollo and Skylab Spacecrafts. The External Tank is 182 feet long, 27 feet in diameter, and carries 1,700,000 pounds of liquid oxygen and liquid hydrogen propellants. The Shuttle System configuration and the interaction of the system elements pose an intricate system sensitivity which must be analyzed to establish the performance and sizing requirements for the External Tank.
A finite-element stress analysis is performed on Helios filament-wound hydrogen tanks to examine the stress field and effect of end dome geometry on the stress field. Each tank is composed of a central circular cylindrical section with either geodesic or hemispherical end domes, which have metallic polar bosses. The tanks are subjected to combined and separate internal pressure and temperature loading conditions, and the stress contributions of each loading component are examined. The tank-wall-polar-boss interfacial meridional tensile stress in the hemispherical dome is found to be approximately 27 percent lower than that in the geodesic dome. The effects of both material anisotropy and the aluminum lining on the intensities of tensile meridional stress at the tank-wall-polar-boss bonding interface are examined.
TDRS-1 was decommissioned on October 28th 2009 following more than 26 years of operation. The Grand Old Dame of the TDRSS constellation wa s launched aboard the maiden voyage of the Space Shuttle Challenger ( STS-6) in April 1983. TDRS-1 survived a malfunction of the Inertial Upper Stage eventually utilizing its own propulsion system to success fully reach its assigned station in geosynchronous orbit. The anomalo us beginning of the TDRS-1 mission was not without lasting consequenc es as the primary reaction control system (A-side) was completely di sabled with an apparent propellant leak and the secondary system (B-s ide) suffered damage to its negative roll thruster rendering the thru ster inoperable. Following decommissioning the challenge to completin g a successful TDRS-1 end-of-mission (EOM) was the implementation of the stringent orbital debris requirements of NPR 8715.6 with a parti ally functioning spacecraft not originally designed to meet those req uirements. The TDRS-1 EOM had three key goals: 1) removal of the spac ecraft from geosynchronous orbit; 2) depletion of the remaining prope llant; and 3) passivation of all other sources of energy storage or generation. The TDRS-1 EOM approach was one of minimizing risks while accomplishing the goals above. The orbit raising portion of EOM was accomplished using deltavelocity operations already proven during pre vious stationchanging maneuvers. The propellant depletion approach wa s necessarily more aggressive as over 20 hours of burn time was requ ired to deplete the remaining fuel. A novel approach utilizing a spin ning, thrusting, passively controlled spacecraft configuration was ut ilized to achieve reasonable burn durations that met schedule constra ints. This nonstandard configuration required careful analysis of ele ctrical, thermal, and communication subsystems. The configuration wa s thoroughly simulated prior to the start of operations and carefully characterized during the initial spin period and first burn. Passiva tion was by definition a unique operation not previously performed wi thin the TDRS 1-7 constellation. Use was made of a TDRS spacecraft si mulator to verify the operational procedure to mitigate risks and pro vide crew training. TDRS-1 orbit raising maneuvers commenced on June 5th 2010 and completed on June 14th with an apogee 370 km and a peri gee 352 km above geosynchronous altitude. 127 kg of propellant were e stimated to be remaining in the tanks at the completion of orbit rais ing. TDRS-1 was placed in its spinning orientation on June 16th and the first fuel depletion burn was performed the following day. A seri es of 10 depletion burns were performed ending on June 26th when both propellant tank pressures experienced dramatic drops. Final passivat ion was performed on June 27 th deactivating electronics, removing t he batteries from the bus and solar arrays, and disabling the space-t o-ground communications equipment. The Second TDRS Ground Terminal (S TGT) continued to open loop track TDRS-1 for several days attempting command reacquisition several times a day. All attempts were unsucce ssful confirming passivation was achieved. The TDRS-1 orbit at passiv ation was an orbital debris compliant 36,319 x 36,128 km in height. W hile differing spacecraft designs may preclude mimicking the exact TD RS-1 EOM approach, the TDRS-1 campaign serves to demonstrate that pr e-NPR 8715.6 designs can be made to meet the requirements resulting i n a reduced orbit debris environment for future missions
Topics covered include: 1) Method to Estimate the Dissolved Air Content in Hydraulic Fluid; 2) Method for Measuring Collimator-Pointing Sensitivity to Temperature Changes; 3) High-Temperature Thermometer Using Cr-Doped GdAlO3 Broadband Luminescence; 4)Metrology Arrangement for Measuring the Positions of Mirrors of a Submillimeter Telescope; 5) On-Wafer S-Parameter Measurements in the 325-508-GHz Band; 6) Reconfigurable Microwave Phase Delay Element for Frequency Reference and Phase-Shifter Applications; 7) High-Speed Isolation Board for Flight Hardware Testing; 8) High-Throughput, Adaptive FFT Architecture for FPGA-Based Spaceborne Data Processors; 9) 3D Orbit Visualization for Earth-Observing Missions; 10) MaROS: Web Visualization of Mars Orbiting and Landed Assets; 11) RAPID: Collaborative Commanding and Monitoring of Lunar Assets; 12) Image Segmentation, Registration, Compression, and Matching; 13) Image Calibration; 14) Rapid ISS Power Availability Simulator; 15) A Method of Strengthening Composite/Metal Joints; 16) Pre-Finishing of SiC for Optical Applications; 17) Optimization of Indium Bump Morphology for Improved Flip Chip Devices; 18) Measuring Moisture Levels in Graphite Epoxy Composite Sandwich Structures; 19) Marshall Convergent Spray Formulation Improvement for High Temperatures; 20) Real-Time Deposition Monitor for Ultrathin Conductive Films; 21) Optimized Li-Ion Electrolytes Containing Triphenyl Phosphate as a Flame-Retardant Additive; 22) Radiation-Resistant Hybrid Lotus Effect for Achieving Photoelectrocatalytic Self-Cleaning Anticontamination Coatings; 23) Improved, Low-Stress Economical Submerged Pipeline; 24) Optical Fiber Array Assemblies for Space Flight on the Lunar Reconnaissance Orbiter; 25) Local Leak Detection and Health Monitoring of Pressurized Tanks; 26) Dielectric Covered Planar Antennas at Submillimeter Wavelengths for Terahertz Imaging; 27) Automated Cryocooler Monitor and Control System; 28) Broadband Achromatic Phase Shifter for a Nulling Interferometer; 29) Super Dwarf Wheat for Growth in Confined Spaces; 30) Fine Guidance Sensing for Coronagraphic Observatories; 31) Single-Antenna Temperature- and Humidity-Sounding Microwave Receiver; 32) Multi-Wavelength, Multi-Beam, and Polarization-Sensitive Laser Transmitter for Surface Mapping; 33) Optical Communications Link to Airborne Transceiver; 34) Ascent Heating Thermal Analysis on Spacecraft Adaptor Fairings; 35) Entanglement in Self-Supervised Dynamics; 36) Prioritized LT Codes; 37) Fast Image Texture Classification Using Decision Trees; 38) Constraint Embedding Technique for Multibody System Dynamics; 39) Improved Systematic Pointing Error Model for the DSN Antennas; 40) Observability and Estimation of Distributed Space Systems via Local Information-Exchange Networks; 41) More-Accurate Model of Flows in Rocket Injectors; 42) In-Orbit Instrument-Pointing Calibration Using the Moon as a Target; 43) Reliability of Ceramic Column Grid Array Interconnect Packages Under Extreme Temperatures; 44) Six Degrees-of-Freedom Ascent Control for Small-Body Touch and Go; and 45) Optical-Path-Difference Linear Mechanism for the Panchromatic Fourier Transform Spectrometer.
An experimental study was initiated that included the long-term testing of the following SOFI materials, which make up the majority of the Thermal Protection System of the Shuttle External Tank: NCFI 24-124 (acreage foam) and BX-265 (close-out foam, including the intertank flange and bipod areas). A potential alternate material, NCFI 27-68 (acreage foam with flame retardant removed), was also tested. Fire chemistry testing was completed on samples that were retrieved after aging/weathering at intervals of 3, 6, and 12 months. The testing included three standard test methods: limiting oxygen index (ASTM G125), radiant panel (ASTM E162), and cone calorimeter (ASTM E1354).
Six large parabolic dish antennas are located at the Goldstone Deep Space Communications Complex north of Barstow, California. As a large-scale facility located in a remote, isolated desert region, the GDSCC operations require numerous on-site storage facilities for gasoline, diesel and hydraulic oil. These essential fluids are stored in underground storage tanks (USTs). Because USTs may develop leaks with the resultant seepage of their hazardous contents into the surrounding soil, local, State and Federal authorities have adopted stringent regulations for the testing and maintenance of USTs. Under the supervision of JPL's Office of Telecommunications and Data Acquisition, a year-long program has brought 27 USTs at the Goldstone Complex into compliance with Federal, State of California and County of San Bernadino regulations. Of these 27 USTs, 15 are operating today, 11 have been temporary closed down, and 1 abandoned in place. In 1989, the 15 USTs now operating at the Goldstone DSCC will be replaced either by modern, double-walled USTs equipped with automatic sensors for leak detection, or by above ground storage tanks. The 11 inactivated USTs are to be excavated, removed and disposed of according to regulation.
An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e., top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.
An experimental Aerodynamic and Aero-Acoustic loads data base was obtained at transonic Mach numbers for the Space Shuttle Launch Vehicle configured with the ASRM Solid Rocket Boosters as an increment to the current flight configuration (RSRB). These data were obtained during transonic wind tunnel tests (IA 613A) conducted in the Arnold Engineering Development Center 16-Foot transonic propulsion wind tunnel from March 27, 1991 through April 12, 1991. This test is the first of a series of two tests covering the Mach range from 0.6 to 3.5. Steady state surface static and fluctuating pressure distributions over the Orbiter, External Tank and Solid Rocket Boosters of the Shuttle Integrated Vehicle were measured. Total Orbiter forces, Wing forces and Elevon hinge moments were directly measured as well from force balances. Two configurations of Solid Rocket Boosters were tested, the Redesigned Solid Rocket Booster (RSRB) and the Advanced Solid Rocket Motor (ASRM). The effects of the position (i.e. top, bottom, top and bottom) of the Integrated Electronics Assembly (IEA) box, mounted on the SRB attach ring, were obtained on the ASRM configured model. These data were obtained with and without Solid Plume Simulators which, when used, matched as close as possible the flight derived pressures on the Orbiter and External Tank base. Data were obtained at Mach numbers ranging from 0.6 to 1.55 at a Unit Reynolds Number of 2.5 million per foot through model angles of attack from -8 to +4 degrees at sideslip angles of 0, +4 and -4 degrees.
During launch of Shuttle Columbia, mission STS-107, a large piece of spray on foam insulation (SOFI) separated from the external tank left bipod ramp area impacting the shuttle orbiter left wing leading edge. "Analysis showed that this large piece of foam struck Columbia on the underside of the left wing after launch. Later, analysis showed that the larger piece struck Columbia on the underside of the left wing, around Reinforced Carbon-Carbon (RCC) panels 5 through 9, at 81.9 seconds after launch. Further photographic analysis revealed that the large foam piece was approximately 21 to 27 inches long and 12 to 18 inches wide and was moving at a relative velocity to the Shuttle stack of 625 to 840 feet per second (416 to 573 miles per hour) at the time of impact." This impact damaged the wing leading edge resulting in loss of orbiter thermal protection. The piece of errant foam was part of a bipod ramp which was designed to meet thermal and aerodynamic requirements in that region of the external tank (ET).
Wind tunnel tests were conducted in a 14- inch wind tunnel with a 0.004 scale model of the space shuttle launch vehicle in order to (1) determine the cause and possible aerodynamic alterations required to eliminate the Orbiter rolling moment couple; (2) determine configuration alterations to alleviate the forward Orbiter external tank loads; and (3) provide data to verify previous data.
The animation depicts NASA's concept for a next-generation Advanced X-ray Timing Mission. The models and their textures doe not necessarily represent the final iteration. Delivery specifications include launch with Taurus II or Falcon 9, mass of 2650 kg, with a circular low earth orbit at approximately 600 km. The inclination depends on the launch vehicle and spacecraft mass. AXTAR's prime instrument will probe the physics of neutron stars and black holes through X-ray timing and spectral measurements. The primary instrument will be the Large Area Timing Array (LATA). The Sky Monitor Clusters configuration consists of 27 Sky Monitor cameras th at are grouped in five clusters. This configuration will achieve approximately 85 percent all sky coverage. Spacecraft components include a science bus to house the LATA of supermodules; a spacecraft bus to house components such as propulsion tanks, avionics, and reaction wheels; solar arrays configured from space-qualified GaAs 3-junction cells; star trackers for attitude knowledge; a propulsion system of four pods, each containing one 100 lbf and two 5 lbf engines; a launch vehicle adaptor; and a radiation shield.
The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) is presented. The IOA effort first completed an analysis of the Electrical Power Generation/Power Reactant Storage and Distribution (EPG/PRSD) subsystem hardware, generating draft failure modes and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The IOA results were then compared to the NASA FMEA/CIL baselines with proposed Post 51-L updates included. A resolution of each discrepancy from the comparison is provided through additional analysis as required. The results of that comparison are documented for the Orbiter EPG/PRSD hardware. The comparison produced agreement on all but 27 FMEAs and 9 CIL items. The discrepancy between the number of IOA findings and NASA FMEAs can be partially explained by the different approaches used by IOA and NASA to group failure modes together to form one FMEA. Also, several IOA items represented inner tank components and ground operations failure modes which were not in the NASA baseline.
NASA is sponsoring the development of a solid oxide electrochemical oxygen separation and compression system. This system is capable of extracting oxygen from a process stream of spacecraft cabin air, and compressing it to >27,000 kPa without mechanical compressors. Process temperatures are >650C, but the materials in contact with the oxygen do not burn. The system is called eCOG-C, Electrochemical Oxygen Generator and Compressor. A system capable of producing high pressure, high purity oxygen may have application as a method of recharging space suit oxygen tanks during human exploration missions. This paper describes the key performance parameters of a space suit oxygen tank recharge system, and places the key performance parameters of eCOG-c in the context of other methods of space suit oxygen tank recharge. This paper compares the eCOG-c configuration to an earlier electrochemical oxygen generator that has been prototyped and tested. Areas of emphasis for eCOG-C development are listed and described.
The NASA Cryogenics Test Laboratory at Kennedy Space Center conducted long-term testing of SOFI materials under actual-use cryogenic conditions. The lab tested NCFI 24-124 (acreage foam), BX-265 (close-out foam, including intertank flange and bipod areas), and a potential alternate material, NCFI 27-68 (acreage foam with the flame retardant removed). Specimens of all three materials were placed at a site that simulated aging (the Vehicle Assembly Building [VAB]) and a site that simulated weathering (Atmospheric Exposure Test Site [beach site]). After aging/ weathering intervals of 3, 6, and 12 months, the samples were retrieved and tested for their ability to absorb moisture under conditions similar to those experienced by the Space Shuttle External Tank (ET) during the loading of cryogenic propellants.
The NTR is a proven technology that generates high thrust and has a specific impulse (Isp (is) approximately 900 s) twice that of today's best chemical rockets. During the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs, twenty rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability - all the requirements needed for a human mission to Mars. Ceramic metal fuel was also evaluated as a backup option. In NASA's recent Mars Design reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program - the 25 klbf 'Pewee' engine is sufficient for a human Mars mission when used in a clustered engine configuration. The 'Copernicus crewed NTR Mars transfer vehicle design developed for DRA 5.0 has significant capability that can enable reusable '1-year' round trip human missions to candidate near Earth asteroids (NEAs) like 1991 JW in 2027, or 2000 SG344 and Apophis in 2028. A robotic precursor mission to 2000 SG344 in late 2023 could provide an attractive Flight Technology Demonstration of a small NTR engine that is scalable to the 25 klbf-class engine used for human missions 5 years later. In addition to the detailed scientific data gathered from on-site inspection, human NEA missions would also provide a valuable 'check out' function for key elements of the NTR transfer vehicle (its propulsion module, TransHab and life support systems, etc.) in a 'deep space' environment prior to undertaking the longer duration Mars orbital and landing missions that would follow. The initial mass in low Earth orbit required for a mission to Apophis is approximately 323 t consisting of the NTR propulsion module ((is) approximately 138 t), the integrated saddle truss and LH2 drop tank assembly ((is) approximately 123 t), and the 6-crew payload element ((is) approximately 62 t). The later includes a multi-mission Space Excursion Vehicle (MMSEV) used for close-up examination and sample gathering. The total burn time and required restarts on the three 25 klbf 'Pewee-class' engines operating at Isp (is) approximately 906 s, are approximately 76.2 minutes and 4, respectively, well below the 2 hours and 27 restarts demonstrated on the NERVA eXperimental Engine, the NRX-XE. The paper examines the benefits, requirements and characteristics of using NTP for the above NEA missions. The impacts on vehicle design of HLV payload volume and lift capability, crew size, and reusability are also quantified.
In the first portion of this work, 1,2,3,3,4,4-hexafluoro-1,2-bis[4-(dimethylhydroxysilyl)phenoxy]cyclobutane and 1,2,3,3,4,4-hexafluoro-1,2-bis[3-(dimethylhydroxysilyl)phenoxy]cyclobutane were prepared and homopolymerized to afford polymers with excellent thermal stability and Tgs of 27 C and -12 C, respectively. Despite the moderately high wt% of fluorin in the polymer structure (23.8%), these polymers had poor fuel resistance. In fact, swelling measurements indicate that these polymers had apparent solubility parameters of about 18.2 J (exp 1/2) m (exp -3/2) (toluene). Copolymerization of the disilanol monomers with fluorosilicone monomers afforded copolymers containing 20-30 wt% of the perfluorocyclobutane-containing structure displayed adequate fuel resistance, enhanced thermal stability, and a Tg low enough to meet the requirements of a High Speed Civil Transport (HSCT) fuel tank sealant. In the second part of this work, trifluorovinylether-terminated oligomers were prepared and polymerized via cyclodimerization. Initially, an alpha, omega-silanol-terminated fluorosilicone was endcapped with trifluorovinylether end groups via a two-step synthetic sequence. The oligomer was thermally cyclodimerized to a polymer that displayed thermal stability similar to that of a fluorosilicone homopolymer. Second, 1,3-bis[4-trifluorovinyl(oxy)phenyl]-1,3-(3,3,3-trifluoropropyl)dimethyldisiloxane and 1,3-bis{3-trifluorovinyl(oxy)phenyl]-1,3-(3,3,3-trifluoropropyl)dimethyldisiloxane were prepared and cyclodimerized to afford polymers that contained pendant trifluoropropyl groups. The pendant trifluoropropyl groups did enhance solvent resistance in aliphatic hydrocarbon solvents, however, no improvement was observed in aromatic hydrocarbon solvents. These polymers also displayed excellent thermal stability. In the last part of this work, a series of monomers was prepared by the DCC-promoted esterification of 4-[trifluorovinyl(oxy)benzoic acid with alpha, omega-functionalized hydrogenated and partially fluorinated alcohols. The monomers were cyclodimerized to the corresponding polymers. The polymers that did not contain beta hydrogens displayed significantly higher thermal stability than the fully hydrogenated polymers. A commercially-available alpha,omega-hydroxy-terminated perfluoropolyether was then functionalized with 4-[trifluorovinyl(oxy)benzoylchloride. An attempt was made to polymerize the resulting oligomer via the cyclodimerization of the terminal trifluorovinylether moieties. Although the viscosity of the oligomer increased significantly during polymerization, Gel Permeation Chromatography (GPC) analysis revealed that the Tetrahydrofuran (THF) soluble portion of the polymer did not have high molecular weight.