Results for “back contact” — Page 17

NASA NTRS2016

Igneous and Sedimentary Compositions from Four Landing Sites on Mars from the Alpha Particle X-Ray Spectrometer (APXS)

The APXS - supported and promoted strongly by Heinrich Waenke - on all four Mars Rovers has returned compositional data from about 1000 rocks and soil targets along the combined traverses of over 60 kilometers. Providing precise and accurate bulk chemistry with typically 16 quantified elements, the APXS is a powerful and versatile tool that when combined with the ability to traverse to key rocks and soils has provided critical information needed to understand the geologic evolution of Mars. APXS data allow comparisons among landing sites, provide ground truth for orbiters and connections back to SNC meteorites. The soils and dust are basaltic in character and represent the average Mars composition similar to Adirondack basalts from Gusev crater but with unambiguous elevated and correlated S, Cl and Zn contents. At all four landing sites the APXS found several rocks with a felsic composition. The similarity is best assessed in a logarithmic ratio plot of rock normalized to the average soil composition (Fig.1). High alkaline, Al, and low Mg, Fe, low S, Cl and Ni, Zn as well as an Fe/Mn ratio of approximately 50 indicate a likely unaltered and igneous origin. Sediments, e.g. the Burns formation, with approximately 25 wt% SO3 at Meridiani Planum have been documented over 10s of kilometers (Fig. 2). This formation is compositionally homogeneous, but showing the removal of MgSO4 and a threefold increase in Cl downhill in 2 craters. The degraded rim of the Noachian crater Endeavour resembles average Mars crust, with local Ca, Mg and Fe sulfate alteration and elevated Mn, some felsic rocks, and high Al, Si and low Fe rocks, possibly indicating clays. Unusual soils at Gusev crater in the area surrounding Home Plate include some very rich in ferric sulfate salts (up to 35 wt% SO3) and some with 90% wt% SiO2, possibly indicating fumerolic activities. Rocks in the Columbia Hills show significant signs of alteration including elevated S, Cl and Br in the abraded interior. At Gale Crater, mudstones with approximately 20% clay, less than 1% SO3 and overall average Mars composition indicate a former habitable environment with low acidity. A remarkable diversity of compositions was found during Curiosity's traverses, which likely stems from material influx from the northern Gale rim. At Pahrump, the base of Mount Sharp, a homogeneous mudstone with lower Mg and Ca was encountered. The composition of this Murray unit can be traced over several kilometers with smooth trends of higher Fe/Mn and lower soluble trace elements Ni and Zn at higher elevation. The likely aeolian Stimson formation with average Mars crustal composition intersects in several places in sharp contact with the Murray formation. Both units show local alteration halos with highly elevated SiO2 of up to 75%, usually correlated with increase in Ti and often P.

Gellert, R.↗

NASA NTRSDate not supplied

Access to Space for NASA Small Sats: Current and Future Needs

Small spacecraft technology advancements have fundamentally shifted how NASA’s Science Mission Directorate (SMD) executes science investigations. To support this approach, the SMD Rideshare Office (SRO) was established in 2020 to lead the definition and implementation of a directorate-wide rideshare strategy. Serving as the central point of contact for coordinating compatible NASA payloads with launch opportunities, the SRO maximizes science, exploration, and technology return on investment by enabling rideshare or other access to space opportunities for small spacecraft on SMD primary mission launches, VADR commercial launch procurements, and other government agency launch opportunities. As NASA seeks to reduce costs and increase the rate of discovery, small satellites and multi manifest access to space have become integral to achieving the agency’s strategic vision. While NASA has created the above-mentioned mechanisms to expand access to space and achieve lower launch costs for its small satellites, many factors have limited full exploit of the opportunity these mechanisms can bring. NASA continues to evolve its mission cultures and technical requirements to adapt and take advantage of burgeoning commercial launch and rideshare advancements. To do so NASA requires collaboration with small satellite manufacturers, principal investigators, and commercial industry partners. Current needs include technical development and design of structurally robust spacecraft buses capable of withstanding varied launch loads, which will increase rideshare interchangeability and versatility. Further, instrument and spacecraft designs must also evolve to handle diverse launch environments and loads factors, while reducing reliance on complex purge and cleanliness constraints, sensitivities to silicones and hydrocarbons, and magnetic requirements. Continued maturation of small and medium launch providers in the near-term is also essential to drive down costs through competition. The current mission selection cadence often complicates the ability to synchronize multiple missions on a single launch. Future needs can include affordable space maneuverability options such as enhanced spacecraft propulsion systems and unique orbital maneuvering capabilities for our individual smallsats or constellations. These emerging capabilities offer a path to unique science orbits for NASA small satellites, but only under the condition that their cost remains affordable and competitive to accommodate inherently smaller mission budgets. Additionally, the projected surge of multiple SMD small satellites launching simultaneously and to unique deep space science orbits necessitates evaluation of expanding deep space communications capabilities. This presentation provides a comprehensive overview of NASA SMD’s access to space landscape and offers further unique insights and discussion, backed by NASA rideshare experiences and lessons learned, on the current and future developments required to unleash the full potential of rideshare opportunities.

Rideshare↗

NASA NTRSDate not supplied

Access to Space for NASA SmallSats: Current and Future Needs

Small spacecraft technology advancements have fundamentally shifted how NASA’s Science Mission Directorate (SMD) executes science investigations. To support this approach, the SMD Rideshare Office (SRO) was established in 2020 to lead the definition and implementation of a directorate-wide rideshare strategy. Serving as the central point of contact for coordinating compatible NASA payloads with launch opportunities, the SRO maximizes science, exploration, and technology return on investment by enabling rideshare or other access to space opportunities for small spacecraft on SMD primary mission launches, VADR commercial launch procurements, and other government agency launch opportunities. As NASA seeks to reduce costs and increase the rate of discovery, small satellites and multi manifest access to space have become integral to achieving the agency’s strategic vision. While NASA has created the above-mentioned mechanisms to expand access to space and achieve lower launch costs for its small satellites, many factors have limited full exploit of the opportunity these mechanisms can bring. NASA continues to evolve its mission cultures and technical requirements to adapt and take advantage of burgeoning commercial launch and rideshare advancements. To do so NASA requires collaboration with small satellite manufacturers, principal investigators, and commercial industry partners. Current needs include technical development and design of structurally robust spacecraft buses capable of withstanding varied launch loads, which will increase rideshare interchangeability and versatility. Further, instrument and spacecraft designs must also evolve to handle diverse launch environments and loads factors, while reducing reliance on complex purge and cleanliness constraints, sensitivities to silicones and hydrocarbons, and magnetic requirements. Continued maturation of small and medium launch providers in the near-term is also essential to drive down costs through competition. The current mission selection cadence often complicates the ability to synchronize multiple missions on a single launch. Future needs can include affordable space maneuverability options such as enhanced spacecraft propulsion systems and unique orbital maneuvering capabilities for our individual smallsats or constellations. These emerging capabilities offer a path to unique science orbits for NASA small satellites, but only under the condition that their cost remains affordable and competitive to accommodate inherently smaller mission budgets. Additionally, the projected surge of multiple SMD small satellites launching simultaneously and to unique deep space science orbits necessitates evaluation of expanding deep space communications capabilities. This presentation provides a comprehensive overview of NASA SMD’s access to space landscape and offers further unique insights and discussion, backed by NASA rideshare experiences and lessons learned, on the current and future developments required to unleash the full potential of rideshare opportunities.

Rideshare↗

NASA NTRSDate not supplied

Error Sources and Mitigation Strategies for Thermocouples Integrated in Flexible Thermal Protection System Materials

Brief Presenter Biography:Ruth Miller is an aer-ospace systems engineer in the Entry Systems and Ve-hicle Development Branch at NASA Ames Research Center.Introduction:The flexible thermal protection sys-tem (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle will be in-strumented with thermocouples (TCs) to measure the in-depth thermal response during entry into Earth’s atmos-phere[1, 2]. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models.However,the deployable nature of inflatable decelerator technology presents challengesfor integrat-ing TCs, specifically the TCsneed to be compactableand cannot damage the FTPSnor the inflatable structure(IS).Unlike traditional rigid aeroshells, routing TCsthrough the thickness of the FTPS could cause signifi-cant damage during packing of the deployable aeroshellbecause the different layers may shift small amounts in relation to each other imparting strain on the TCsand FTPS materials. The LOFTID TCleads are routed from the measurement location back to the data acquisition system in the vehicle centerbody within the same FTPS layer that they are monitoring the temperature. This ap-proach eliminates the need to put holes in the FTPS lay-ers,butas a consequence,the insulated TCleads travel for an appreciable distance through a region that will expose them to high temperaturesand large thermal gra-dients.LOFTID’s TCswere baselined to be commercially available Type K TCswith a binder impregnated glass braid insulation. These TCswere chosen because they had been used successfully on IRVE-3 and in ground-based arc jet testing. Additionally, these TCsdid notdamage the FTPS nor IS during packing and deploy-ment testing. However, the glass braid insulation is only rated to a maximum continuous use temperature of 482ºC. For reference, the duration of the heat pulse on the LOFTID vehicle is on the order of minutes and themaximum predicted temperaturebeneaththe outermost FTPS layersis 1350ºC.Ground-based testing in a tube furnace at NASA Ames Research Centerwas conducted to determine if the baseline TCsrouted through FTPS samples would survive and provide accurate temperature measure-mentsat LOFTID flight-relevant temperatures[3].The test results showed large measurement errors occurred beginning at approximately 400°C due to conductive deposits on the TCinsulation electrically shorting the TCleads.The conductive deposits and thus electrical shorting weredetermined to be caused by twoerror sources:1.The organic binder on the TCinsulation carbon-izingin a high temperature, low oxygen envi-ronment2.Decomposition products from the FTPSperme-atingthe braided TC insulationFurther testing in the tube furnace demonstrated that heat cleaning the TCinsulation effectivelyremovedthe organic binder andeliminatedthe first error source.The second error sourcewas shown to be mitigated by the addition of amica wraparound each individual TCleadtoact as an impermeable barrier.To understand the applicability of theground-based tube furnace test results to flight,anarc jet testseriesat Boeing’s Large Core Arc Tunnel (LCAT) facilitywas conducted[4].The error sourcesand mitigation strate-giesidentified in the tube furnace testing were substan-tiatedin the arc jet testing.However, the arc jet testing also revealed three new error sources:1.Glass TCinsulation meltingwhich resultsin electrical shorting of the TCeither through di-rect contact between the two leads or through the electrically conductive FTPS materials2.TCwire meltingwhich results in a noisy and/or open-loop TCresponse3.Type K TCwire green-rotwhich results in large calibration errors Scope of the Presentation:This presentation will include a brief discussion onthe effect of electrical shorting on the output of a TC(i.e. how to identify elec-trical shorting in TCdataand what the associated erroris).The tube furnace and arc jet test resultswill be dis-cussedand the solutions LOFTID is implementing to mitigate the error sourcesidentified in the tube furnace and arc jet testingwill be presented.Additionally, futureresearch and development work to eliminateTCerror sourcesfor future missionswill be recommended.

R A Miller↗

NASA NTRS2016

Modifying a Commercial Centrifuge to Reduce Electromagnetic Interference and Evaluating Functionality of Ultrasound Equipment

The Project Management and Engineering Branch (SF4) supports the Human Health and Performance Directorate (HH&P) and is responsible for developing and supporting human systems hardware for the International Space Station (ISS). When a principal investigator's (PI) medical research project on the ISS is accepted, SF4 develops the necessary hardware and software to transport to the ISS. The two projects I primarily worked on were the centrifuge and ultrasound projects. Centrifuge: One concern with spacecraft such as the ISS is electromagnetic interference (EMI) from onboard equipment, typically from radio waves (frequencies of ~3 kHz to ~300 GHz), which can negatively affect nearby circuitry. Standard commercial centrifuges produce EMI above safety limits, so my task was to help reduce EMI production from this equipment. Two centrifuges were tested: one unmodified as a control and one modified. To reduce EMI below safety limits, one centrifuge was modified to become a Faraday shield, in which significant electrical contact was made between all regions of the centrifuge housing. This included removing non-conductive paint, applying conductive fabric to the lid and foam sealer, adding a 10,000 μF decoupling capacitor across the power supply, and adding copper adhesive-mount gaskets to the housing interior. EMI testing of both centrifuges was performed in the EMI/EMC Control Test and Measurement Facility. EMI for both centrifuges was below safety limits for frequencies between 10 MHz and 15 GHz (pass); however, between 14 kHz and 10 MHz, EMI for the unmodified centrifuge exceeded safety limits (fail) as expected. Alternatively, for the modified centrifuge with the Faraday shield, EMI was below the safely limit of 55 dBμV/m for electromagnetic frequencies between 14 kHz and 10 MHz. This result indicates our modifications were successful. The successful EMI test allowed us to communicate with the vendor what modifications they needed to make to their commercial unit to meet our specifications and to understand what needs to be done in lab to the new centrifuge. Our modifications will provide a standard for readying centrifuges for future missions. Once the new modified centrifuge arrives by the vendor, it will need to undergo EMI testing again for validation. The centrifuge is also in the process of compatibility testing with a custom stowage drawer, which is an ongoing project in SF4. Both of these items will be payloads on future missions to the ISS for various research purposes. Ultrasound: ISS currently has an onboard ultrasound (Ultrasound 2 system) for research and medical purposes. Every piece of medical flight hardware has an equivalent ground-unit so instrumentation can be routinely evaluated and transported to the ISS if necessary. The ground-unit ultrasound equipment must be evaluated every six months using a task performance sheet (TPS). A TPS is a document, written by the appropriate scientists and engineers, which describes how to run equipment and is written in such a way that astronauts with unspecialized training can follow the tasks. I was responsible for performing six TPSs on a combination of three ultrasounds and two video power converters (VPCs). Performing a TPS involves checking out and computationally documenting each piece of equipment removed from storage locations, setting up hardware and software, performing tasks to verify functionality, returning equipment, and logging items back into the computerized system. My work revealed all ground-unit ultrasounds were functioning properly. Because of proper function, a discrepancy report (DR) did not have to be opened. The TPS was then passed along to the Quality Engineering (QE) for review and ultimately given to Quality Assurance (QA). Other projects: In addition to my main projects, I participated in other tasks including troubleshooting an EEG headband, volunteering for an ultrasound training research study, and conformal coating printed circuit boards. My internship at SF4 has helped me understand how space systems hardware development for the ISS fits into NASA's mission and vision.

Greening, Gage J.↗

NASA NTRS2015

Development of Long-Lifetime Pulsed Gas Valves for Pulsed Electric Thrusters

It is advantageous for gas-fed pulsed electric thrusters to employ pulsed valves so propellant is only flowing to the device during operation. The propellant utilization of the thruster will be maximized when all the gas injected into the thruster is acted upon by the fields produced by the electrical pulse. Gas that is injected too early will diffuse away from the thruster before the electrical pulse can act to accelerate the propellant. Gas that is injected too late will miss being accelerated by the already-completed electrical pulse. As a consequence, the valve must open quickly and close equally quickly, only remaining open for a short duration. In addition, the valve must have only a small amount of volume between the sealing body and the thruster so the front and back ends of the pulse are as coincident as possible with the valve cycling, with very little latent propellant remaining in the feed lines after the valve is closed. For a real mission of interest, a pulsed thruster can be expected to pulse at least 10(exp 10) - 10(exp 11) times, setting the range for the number of times a valve must open and close. The valves described in this paper have been fabricated and tested for operation in an inductive pulsed plasma thruster (IPPT) for in-space propulsion. In general, an IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged, producing a high-current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed, it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. The valve characteristics needed for the IPPT application require a fast-acting valve capable of a minimum of 10(exp 10) valve actuation cycles. Since even 10(exp 9) cycles is well above anything demonstrated, this lower value was selected as the design point for the present work. The valve seal must remain leak-tight throughout operation, and the body must maintain a low internal leakage at relatively high operating temperatures. The full set of design requirements used for this program are summarized in Table 1. In this work, we describe two pulsed gas valves that have been fabricated to have long lifetime and demonstrate the characteristics listed above. The first is a miniaturized, conventional electromagnet-based valve while the second is a piezoelectric-based valve design. The conventional valve, shown in Fig. 1, is opened by use of a solenoid electromagnetic actuator. When current is applied to the solenoid coil, magnetic forces pull the plunger away from the valve seat, allowing fluid to flow through the valve. Removal of electrical current permits the spring and fluid pressure to seat the plunger, halting the flow of fluid. The valve body is fabricated from 304L corrosion resistant steel (CRES) and while the parts that form the magnetic circuit are fabricated from 430 CRES. This material does not have optimum magnetic properties, but its corrosion resistance permits incorporation into a design without requiring an additional plating process. A viton O-ring compound (Parker V0884-75), selected for its mechanical strength at elevated temperatures, was used for the valve seat seal. The design was based solely on the use of analytical sizing calculations, as opposed to a more rigorous finite element analysis. While this valve is small and relatively lightweight, it does not represent a design that is optimized for mass and/or a given volume envelope. The piezoelectric valve is a "puller" valve design. Applying a voltage to the piezo crystal causes it to elongate and pull a pintle off the seat, opening the valve. The valve seal consists of the pintle with an external, spherically-formed tip fabricated from Inconel 625 and the seat fabricated from Inconel 718 and possessing a spherical surface that permits the pintle to align with the valve body centerline after each cycle, maintaining the circumferential metal-on-metal seal. The contact stress on the valve pintle is expected in the worst case to be 96.5 MPa (14 ksi), which is well below the work hardening value of 413.7 MPa (60 ksi) for Inconel 615. Results presented in this work will compare the performance of each valve, both to each other and to the valve design requirements given in Table 1. In general, both valves meet all the performance envelope requirements, but the piezoelectric valve consumes considerably less power and has a faster response time over the conventional solenoid-driven valves.

Burkhardt, Wendel M.↗