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Biosatellite Primate Mission Program for component qualification levels validated by flight tests data, performing vibration tests on prime structure spacecraft
Current practices used to design space payload structures for transient dynamic loads are described. Included in the discussion of the design process are the estimation of preliminary loads, development of dynamic models, calculation of transient loads, verification of analytical models, establishment of development tests, definition of the qualification tests, and acquisition of flight data. An indication of the accuracy of the analytical predictions is provided by comparing them with actual test and flight data. Many illustrations and much data obtained during the Viking Project are used to clarify the current practices.
A study has been completed to define a Shuttle experiment that solves the most crucial scientific and engineering problems involved in building a satellite Doppler wind profiler for making global wind measurements. The study includes: (1) a laser study to determine the feasibility of using the existing NOAA Windvan laser in the Space Shuttle spacecraft; (2) a preliminary optics and telescope design; (3) an accommodations study including power, weight, thermal, and control system requirements; and (4) a flight trajectory and operations plan designed to accomplish the required scientific and engineering goals. The experiment will provide much-needed data on the global distribution of atmospheric aerosols and demonstrate the technique of making wind measurements from space, including scanning the laser beam and interpreting the data. Engineering accomplishments will include space qualification of the laser, development of signal processing and lag angle compensation hardware and software, and telescope and optics design. All of the results of this limited Spacelab experiment will be directly applicable to a complete satellite wind profiler for the Earth Observation System/Space Station or other free-flying satellite.
The US National Oceanic and Atmospheric Administration (NOAA) operates the Polar Operational Environmental Satellite (POES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the POES series of spacecraft, named as NOAA-KLMNN', one is in orbit and four more are in various phases of development. The NOAA-K spacecraft was launched on May 13, 1998. Each of these spacecraft carry three Nickel-Cadmium batteries designed and manufactured by Lockheed Martin. The battery, which consists of seventeen 40 Ah cells manufactured by SAFT, provides the spacecraft power during the ascent phase, orbital eclipse and when the power demand is in excess of the solar array capability. The NOAA-K satellite is in a 98 degree inclination, 7:30AM ascending node orbit. In this orbit the satellite experiences earth occultation only 25% of the year. This paper provides a brief overview of the power subsystem, followed by the battery design and qualification, the cell life cycle test data, and the performance during launch and in orbit.
The US National Oceanic and Atmospheric Administration (NOAA) operates the Polar Operational Environmental Satellite (POES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the POES series of spacecraft, named as NOAA-KLMNN, is in orbit and four more are in various phases of development. The NOAA-K spacecraft was launched on May 13, 1998. Each of these spacecraft carry three Nickel-Cadmium batteries designed and manufactured by Lockheed Martin. The battery, which consists of seventeen 40 Ah cells manufactured by SAFT, provides the spacecraft power during the ascent phase, orbital eclipse and when the power demand is in excess of the solar array capability. The NOAA-K satellite is in a 98 degree inclination, 7:30AM ascending node orbit. In this orbit the satellite experiences earth occultation only 25% of the year. This paper provides a brief overview of the power subsystem, followed by the battery design and qualification, the cell life cycle test data, and the performance during launch and in orbit.
The Propulsive Small Expendable Deployer System (ProSEDS) mission is designed to provide an on-orbit demonstration of the electrodynamic propulsion capabilities of tethers in space. The ProSEDS experiment will be a secondary payload on a Delta 11 unmanned expendable booster. A 5-km conductive tether is attached to the Delta 11 second stage and collects current from the low Earth orbit (LEO) plasma. A hollow cathode plasma contactor emits the collected electrons from the Delta II, completing the electrical circuit with the ambient plasma. The current flowing through the tether generates thrust based on the Lorentz Force Law. The thrust will be generated opposite to the velocity vector, slowing down the spacecraft and causing it to de-orbit in approximately 14 days compared to the normal 6 months. A 10-km non-conductive tether is between the conductive tether and an endmass containing several scientific instruments. The ProSEDS mission lifetime was set at I day because most of the primary objectives can be met in that time. The extended ProSEDS mission will be for as many days as possible, until the Delta 11 second stage burns up or the tether is severed by a micrometeoroid or space debris particle. The Hollow Cathode Plasma Contactor (HCPC) unit has been designed for a 12-day mission. Because of the science requirements to measure the background ambient plasma, the HCPC must operate on a duty cycle. Later in the ProSEDS mission, the HCPC is operated in a manner to allow charging of the secondary battery. Due to the unusual operating requirements by the ProSEDS mission, a development unit of the HCPC was built for thorough testing. This developmental unit was tested for a simulated ProSEDS mission, with measurements of the ability to start and stop during the duty cycle. These tests also provided valuable data for the ProSEDS software requirements. Qualification tests of the HCPC flight hardware are also discussed.
Within the pressurized elements of the International Space Station (ISS), requirements exist to ensure a safe, habitable environment for the crew. In order to provide this environment, thermal control components work in conjunction with software controls to provide heat rejection for subsystem avionics equipment, for the environmental control system and for experiment payloads. It is essential to ISS operations, mission success and crew safety that necessary testing incorporates the extreme conditions to ensure proper performance. This paper provides a general description and methodology applied to thermal related Hardware/Software Integration (HSI) tests for the ISS Node 2 module. A detailed test plan was developed and implemented with two objectives: the first was for risk mitigation of the thermal control algorithms and software qualification, and the second was for data collection which will substantiate thermalhydraulic models of the Internal Active Thermal Control System (IATCS). Analytical models are utilized to determine on-orbit performance for conditions and scenarios where the simulation of actual on-orbit system performance is limited by test configuration constraints. Node 2 IATCS HSI activities were performed at the Alenia Spazio facility in Torino, Italy with participation from the National Aeronautics and Space Administration (NASA), Alenia Spazio, Jacobs Engineering Sverdrup (JE Sverdrup) and Boeing.
Radiation Hardness Assurance (RHA) methodologies against Total Ionizing Dose (TID) degradation impose rigorous statistical treatments for data from a part's Radiation Lot Acceptance Test (RLAT) and/or its historical performance. However, no similar methods exist for using "similarity" data - that is, data for similar parts fabricated in the same process as the part under qualification. This is despite the greater difficulty and potential risk in interpreting of similarity data. In this work, we develop methods to disentangle part-to-part, lot-to-lot and part-type-to-part-type variation. The methods we develop apply not just for qualification decisions, but also for quality control and detection of process changes and other "out-of-family" behavior. We begin by discussing the data used in ·the study and the challenges of developing a statistic providing a meaningful measure of degradation across multiple part types, each with its own performance specifications. We then develop analysis techniques and apply them to the different data sets.
A compressor face rake with an internal valve arrangement to permit nulling was designed, constructed, and tested in the laboratory and in flight at the NASA Flight Research Center. When actuated by the pilot in flight, the nullable rake allowed the transducer zero shifts to be determined and then subsequently removed during data reduction. Design details, the fabrication technique, the principle of operation, brief descriptions of associated digital zero-correction programs and the qualification tests, and test results are included. Sample flight data show that the zero shifts were large and unpredictable but could be measured in flight with the rake. The rake functioned reliably and as expected during 25 hours of operation under flight environmental conditions and temperatures from 230 K (-46 F) to greater than 430 K (314 F). The rake was nulled approximately 1000 times. The in-flight zero-shift measurement technique, as well as the rake design, was successful and should be useful in future applications, particularly where accurate measurements of both steady-state and dynamic pressures are required under adverse environmental conditions.
The NASA Protoflight Research Initiative is an internal NASA study conducted within the Office of the Chief Engineer to better understand the use of Protoflight within NASA. Extensive literature reviews and interviews with key NASA members with experience in both robotic and human spaceflight missions has resulted in three main conclusions and two observations. The first conclusion is that NASA's Protoflight method is not considered to be "prescriptive." The current policies and guidance allows each Program/Project to tailor the Protoflight approach to better meet their needs, goals and objectives. Second, Risk Management plays a key role in implementation of the Protoflight approach. Any deviations from full qualification will be based on the level of acceptable risk with guidance found in NPR 8705.4. Finally, over the past decade (2004 - 2014) only 6% of NASA's Protoflight missions and 6% of NASA's Full qualification missions experienced a publicly disclosed mission failure. In other words, the data indicates that the Protoflight approach, in and of it itself, does not increase the mission risk of in-flight failure. The first observation is that it would be beneficial to document the decision making process on the implementation and use of Protoflight. The second observation is that If a Project/Program chooses to use the Protoflight approach with relevant heritage, it is extremely important that the Program/Project Manager ensures that the current project's requirements falls within the heritage design, component, instrument and/or subsystem's requirements for both the planned and operational use, and that the documentation of the relevant heritage is comprehensive, sufficient and the decision well documented. To further benefit/inform this study, a recommendation to perform a deep dive into 30 missions with accessible data on their testing/verification methodology and decision process to research the differences between Protoflight and Full Qualification missions' Design Requirements and Verification & Validation (V&V) (without any impact or special request directly to the project).
The cryogenics analysis program was developed as a simplified tool for use in premission planning operations for the Apollo command service module. Through a dynamic development effort, the program has been extended to include real time and postflight analysis capabilities with nominal and contingency planning features. The technical aspects of the program and a comparison of ground test and mission data with data generated by using the cryogenics analysis program are presented. The results of the program capability to predict flight requirements also are presented. Comparisons of data from the program with data from flight results, from a tank qualifications program, and from various system anomalies that have been encountered are discussed. Future plans and additional considerations for the program also are included. Among these plans are a three tank management scheme for hydrogen, venting profile generation for Skylab, and a capability for handling two gas atmospheres. The plan for two gas atmospheres will involve the addition of the capability to handle nitrogen as well as oxygen and hydrogen.
Mortars are used as one method for ejecting parachutes into the airstream to decelerate spacecraft and aircraft pilot escape modules and to effect spin recovery of the aircraft. An approach to design of mortars in the class that can accommodate parachutes in the 20- to 55-foot-diameter size is presented. Parachute deployment considerations are discussed. Comments are made on the design of a power unit, mortar tube, cover, and sabot. Propellant selection and breech characteristics and size are discussed. A method of estimating hardware weights and reaction load is presented. In addition, some aspects of erodible orifices are given as well as comments concerning ambient effects on performance. This paper collates data and experience from design and flight qualification of four mortar systems, and provides pertinent estimations that should be of interest on programs considering parachute deployment.
Established are the requirements for performance, design, test and qualification of one type of equipment identified as SMS C&D panels and forward structures. This CEI is used to provide all hardware and wiring necessary for the C&D panels to be properly interfaced with the computer complex/signal conversion equipment (SCE), crew station, and software requirements as defined in other CEI specifications.
One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster's operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry, which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.
One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster s operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. There are many different sensors and techniques that have been employed to quantify discharge channel erosion in Hall thrusters. Snapshots of the wear pattern can be obtained at regular shutdown intervals using laser profilometry. Many non-intrusive techniques of varying complexity and sensitivity have been employed to detect the time-varying presence of erosion products in the thruster plume. These include the use quartz crystal microbalances, emission spectroscopy, laser induced flourescence, and cavity ring-down spectroscopy. While these techniques can provide a very accurate picture of the level of eroded material in the thruster plume, it is more difficult to use them to determine the location from which the material was eroded. Furthermore, none of the methods cited provide a true in-situ measure of erosion at the channel surface while the thruster is in operation (i.e. none yield a continuous channel erosion measurement). A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry (see Fig. 1), which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.
Although oxygen is a chemically stable element, it is not shock sensitive, will not decompose, and is not flammable. Oxygen use therefore carries a risk that should never be overlooked, because oxygen is a strong oxidizer that vigorously supports combustion. Safety is of primary concern in oxygen service. To promote safety in oxygen systems, the flammability of materials used in them should be analyzed. At the NASA White Sands Test Facility (WSTF), we have performed configurational tests of components specifically engineered for oxygen service. These tests follow a detailed WSTF oxygen hazards analysis. The stated objective of the tests was to provide performance test data for customer use as part of a qualification plan for a particular component in a particular configuration, and under worst-case conditions. In this document - the 'Guide for Oxygen Component Qualification Tests' - we outline recommended test systems, and cleaning, handling, and test procedures that address worst-case conditions. It should be noted that test results apply specifically to: manual valves, remotely operated valves, check valves, relief valves, filters, regulators, flexible hoses, and intensifiers. Component systems are not covered.
This document contains the procedure and the test results of the Advanced Microwave Sounding Unit-A (AMSU-A) Electromagnetic Interference (EMI), Electromagnetic Susceptibility, and Electromagnetic Compatibility (EMC) qualification test for the Meteorological Satellite (METSAT) and the Meteorological Operation Platform (METOP) projects. The test was conducted in accordance with the approved EMI/EMC Test Plan/Procedure, Specification number AE-26151/5D. This document describes the EMI/EMC test performed by Aerojet and it is presented in the following manner: Section-1 contains introductory material and a brief summary of the test results. Section 2 contains more detailed descriptions of the test plan, test procedure, and test results for each type of EMI/EMC test conducted. Section 3 contains supplementary information that includes test data sheets, plots, and calculations collected during the qualification testing.
NASA's future missions are focused on deep space for human exploration that do not provide a simple emergency return to Earth. In addition, the deep space environment contains a constant background Galactic Cosmic Ray (GCR) radiation exposure, as well as periodic Solar Particle Events (SPEs) that can produce intense amounts of radiation in a short amount of time. Given these conditions, it is important that the avionics systems for deep space human missions are not susceptible to Single Event Effects (SEE) that can occur from radiation interactions with electronic components. The typical process to minimizing SEE effects is through using heritage hardware and extensive testing programs that are very costly. Previous work by Koontz, et al. [1] utilized an analysis-based method for investigating electronic component susceptibility. In their paper, FLUKA, a Monte Carlo transport code, was used to calculate SEE and single event upset (SEU) rates. This code was then validated against in-flight data. In addition, CREME-96, a deterministic code, was also compared with FLUKA and in-flight data. However, FLUKA has a long run-time (on the order of days), and CREME-96 has not been updated in several years. This paper will investigate the use of HZETRN 2010, a deterministic transport code developed at NASA Langley Research Center, as another tool that can be used to analyze SEE and SEU rates. The benefits to using HZETRN over FLUKA and CREME-96 are that it has a very fast run time (on the order of minutes) and has been shown to be of similar accuracy as other deterministic and Monte Carlo codes when considering dose [2, 3, 4]. The 2010 version of HZETRN has updated its treatment of secondary neutrons and thus has improved its accuracy over previous versions. In this paper, the Linear Energy Transfer (LET) spectra are of interest rather than the total ionizing dose. Therefore, the LET spectra output from HZETRN 2010 will be compared with the FLUKA and in-flight data to validate HZETRN 2010 as a computational tool for SEE qualification by analysis. Furthermore, extrapolation of these data to interplanetary environments at 1 AU will be investigated to determine whether HZETRN 2010 can be used successfully and confidently for deep space mission analyses.