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The opening, closing, and latching of the first large clamshell door to operate in space presented some unusual challenges for the mechanism designer of the Space Shuttle Program. The requirements, hardware configuration, design trade-offs, and qualification testing in process to meet the challenge and to make the system operational for the Shuttle orbiter's approaching first orbital flight are described.
Program plan for earth orbital low gravity heat transfer and fluid mechanics experiments
The current version of the NASA turbulent simulation model in the form of a digital computer program, TBMOD, is described. The logic of the program is discussed and all inputs and outputs are defined. An alternate method of shear simulation suitable for incorporation into the model is presented. The simulation is based on a von Karman spectrum and the assumption of isotropy. The resulting spectral density functions for the shear model are included.
One of four major areas of project Pathfinder is in-space assembly and construction. The task of in-space assembly and construction is to develop the requirements and the technology needed to build elements in space. A 120-ft diameter tetrahedral aerobrake truss is identified as the focus element. A heavily loaded mechanical joint is designed to robotically assemble the defined aerobrake element. Also, typical large components such as habitation modules, storage tanks, etc., are defined, and attachment concepts of these components to the tetrahedral truss are developed.
Work performed by Southwest Research Institute (SwRI) as part of an Independent Technical Assessment (ITA) for the NASA Engineering and Safety Center (NESC) is summarized. The ITA goal was to establish a flight rationale in light of a history of fatigue cracking due to flow induced vibrations in the feedline flowliners that supply liquid hydrogen to the space shuttle main engines. Prior deterministic analyses using worst-case assumptions predicted failure in a single flight. The current work formulated statistical models for dynamic loading and cryogenic fatigue crack growth properties, instead of using worst-case assumptions. Weight function solutions for bivariant stressing were developed to determine accurate crack "driving-forces". Monte Carlo simulations showed that low flowliner probabilities of failure (POF = 0.001 to 0.0001) are achievable, provided pre-flight inspections for cracks are performed with adequate probability of detection (POD)-specifically, 20/75 mils with 50%/99% POD. Measurements to confirm assumed POD curves are recommended. Since the computed POFs are very sensitive to the cyclic loads/stresses and the analysis of strain gage data revealed inconsistencies with the previous assumption of a single dominant vibrant mode, further work to reconcile this difference is recommended. It is possible that the unaccounted vibrational modes in the flight spectra could increase the computed POFs.
We present the findings of the test program performed by The European Space Tribology Laboratory (ESTL) to evaluate the performance (friction and lifetime) of a number of space lubricants under vacuum using a Spiral Orbit Tribometer (SOT). Focus was given to a comparison of various popular space oils, a comparison study between the old and new MAPLUB grease formulations, and the performance of commonly used solid lubricants under various conditions. Tests demonstrated that the lifetimes of hydrocarbon NYE oils 2001 & 2001A outperformed those of the perfluroropolyalkylether (PFPE) oils Fomblin Z25 & Z60, though these pairs displayed similar behavior. This relationship was also generally seen for greases; with the lifetimes of the multiple alkylated cyclopentane (MAC)-based greases being extended in comparison to the PFPE-based greases. Testing on greases also demonstrated similar performance between the old (-a) and new (-b) formulations when considering PFPE-based MAPLUB greases, and indeed for all tested PFPE-based non-MAPLUB greases, but significantly shorter lifetimes for the new formulations when considering MAC-based MAPLUB greases. MAPLUB MAC greases containing molybdenum disulphide (MoS2) thickener were also found to display reduced lifetimes. For solid lubricants, lead displayed significantly extended lifetimes over MoS2, speculated to be caused by redistribution of lead from the ball onto all contact surfaces during the test. Friction coefficients were seen to be some 2.5x higher for lead than for MoS2 under similar conditions, a result that corresponds well with conventional bearing tests. The work described was performed under contract for the European Space Agency as part of the Tribology Applications Program, with all funding for testing and apparatus provided by European Space Agency (ESA).
Exploration of the Ice Giants, especially Uranus, via orbiter and atmospheric probes, is required to answer pressing science questions that have been raised in the latest National Academies of Sciences Planetary Decadal Survey. Since the Ice Giants are the farthest planets from Earth, traditional fully-propulsive orbit insertion missions have transit times to the planetary bodies bordering 13-15 years and require a large amount of propellant (wet mass percentages of around 60-70%) for the orbit insertion maneuver, leaving less mass for the scientific payload and a planetary probe. Aerocapture uses aerodynamic forces generated by flight within a planetary atmosphere to decelerate and achieve orbit insertion. Aerocapture has been considered for several past missions but it has not been demonstrated. However, recent developments in thermal protection systems (TPS), guidance and control (G&C), and interplanetary navigation capabilities show the potential for using rigid, heritage entry vehicle configurations already flown at other planetary bodies for Ice Giants aerocapture. Aerocapture can robustly deliver spacecraft to Ice Giant orbits, while substantially increasing on-orbit payload mass (more than 40%) that can be used for a robust atmospheric entry probe. Additionally, the aerocapture maneuver would reduce the interplanetary transit time by 2-5 years (15-30%) relative to fully-propulsive orbit insertion. Recent work has shown that a flagship-class mission can be conducted in a shorter time than fully-propulsive missions if using aerocapture. This paper will consider the merits of including aerocapture as the orbit-insertion mechanism for a Uranus mission. Specifically, the implications of aerocapture orbit insertion for in-situ atmospheric probes will be discussed. The Uranus Orbiter and Probe concept mission study [3] is considered as the potential payload. Results from a recent NASA Space Technology Mission Directorate (STMD)-funded activity that is designing an aerocapture mission for a Uranus orbiter will be presented.
Exploration of the Ice Giants, especially Uranus, via orbiter and atmospheric probes, is required to answer pressing science questions that have been raised in the latest National Academies of Sciences Planetary Decadal Survey. Since the Ice Giants are the farthest planets from Earth, traditional fully-propulsive orbit insertion missions have transit times to the planetary bodies bordering 13-15 years and require a large amount of propellant (wet mass percentages of around 60-70%) for the orbit insertion maneuver, leaving less mass for the scientific payload and a planetary probe. Aerocapture uses aerodynamic forces generated by flight within a planetary atmosphere to decelerate and achieve orbit insertion. Aerocapture has been considered for several past missions but it has not been demonstrated. However, recent developments in thermal protection systems (TPS), guidance and control (G&C), and interplanetary navigation capabilities show the potential for using rigid, heritage entry vehicle configurations already flown at other planetary bodies for Ice Giants aerocapture. Aerocapture can robustly deliver spacecraft to Ice Giant orbits, while substantially increasing on-orbit payload mass (more than 40%) that can be used for a robust atmospheric entry probe. Additionally, the aerocapture maneuver would reduce the interplanetary transit time by 2-5 years (15-30%) relative to fully-propulsive orbit insertion. Recent work has shown that a flagship-class mission can be conducted in a shorter time than fully-propulsive missions if using aerocapture. This paper will consider the merits of including aerocapture as the orbit-insertion mechanism for a Uranus mission. Specifically, the implications of aerocapture orbit insertion for in-situ atmospheric probes will be discussed. The Uranus Orbiter and Probe concept mission study [3] is considered as the potential payload. Results from a recent NASA Space Technology Mission Directorate (STMD)-funded activity that is designing an aerocapture mission for a Uranus orbiter will be presented.
The Inertial Upper Stage (IUS) is a two-stage booster used by NASA and the Defense Department to insert payloads into geosynchronous orbit from low-Earth orbit. The hydraulic mechanism discussed here was designed to perform a specific dynamic and static interface function within the Space Transportation System's Orbiter. Requirements, configuration, and application of the hydraulic mechanism with emphasis on performance and methods of achieving zero external hydraulic leakage are discussed. The hydraulic load-leveler mechanism meets the established design requirements for operation in a low-Earth orbit. Considerable testing was conducted to demonstrate system performance and verification that external leakage had been reduced to zero. Following each flight use of an ASE, all hydraulic mechanism components are carefully inspected for leakage. The ASE, including the hydraulic mechanism, has performed without any anomalies during all IUS flights.
Voyager photographic data are used for microscopic and macroscopic analyses of the dynamics of the Saturn rings. Attention is given to particle collisions, local velocity dispersion, the particle size distribution, mechanisms governing the ring structure and interactive effects with Saturn satellite orbits. The mechanical properties and collision dynamics of the particles are discussed, along with models for perturbed and unperturbed disks. The forces involved and true scale of particle erosion are considered, as are instabilities in the rings, satellite torques, resonances, and the time scale for the evolution of the satellite orbits. Finally, the possibility that angular momentum exchanges are occurring between individual rings and satellite orbits is examined.
Scanning Equipment supporting the Millimeter Wave Radiometer Instrument (MWRI) are flying in a sunsynchronized orbit of 850-km altitude with an inclination of 98.8 deg on the FY-3 meteorological satellite (FY = Feng Yun, Wind and Cloud). MWRI is a linearly polarized, ten-channel passive Radiometer; it measures precipitation and water clouds, sea ice, snow/water equivalent, drought and flood index, land temperature and soil moisture. Following the FY3-A, the FY3-B Satellite was launched in autumn 2010. Since that time, the Scanning Equipment was continuously operated. During the last three and a half years in orbit, the Scanning Mechanism has executed about 65 million revolutions, while the Scan Compensation Mechanism (SCM) - used for momentum compensation - has already successfully executed more than one billion revolutions. During the commissioning phase of the instrument and during the first operation phase, random torque spikes, which manifested themselves as a motor current increase, were observed in the Scan Drive Mechanism, whereas the Scan Compensation drive operated nominally from the beginning. The result of the root cause investigations performed in order to isolate the issue, and the consequences for the follow-on MWRI equipment which was successfully launched by end of September 2013 (now flying on the FY 3-C Spacecraft), are discussed.
Liquid hydrogen fluid mechanics low-gravity problems of Saturn V/S-IVB and control measures
Investigation of new orbit geometries exhibits a very attractive behavior for a spacecraft to monitor space weather coming from the Sun. Several orbit transfer mechanisms are analyzed as potential alternatives to monitor solar activity such as a sub-solar orbit or quasi-satellite orbit and short and long heteroclinic and homoclinic connections between the triangular points L(sub 4) and L(sub 5) and the collinear point L(sub 3) of the Circular Restricted Three-Body Problem (CRTBP) in the Sun-Earth system.
A first-order perturbation theory exists for the flyby problem with the Born approximation as the reference orbit.
Saturn V/S-IVB stage low gravity fluid mechanics problems and Saturn IB liquid hydrogen orbital experiment to verify propellant ullaging and orbital venting system
In May 2003, a new and exciting chapter in Mars exploration will begin with the launch of the first of three spacecraft that will collectively contribute toward the goal of delivering samples from the Red Planet to Earth. This mission is called Mars Sample Return (MSR) and will utilize both the 2003 and 2005 launch opportunities with an expected sample return in October 2008. NASA and CNES are major partners in this mission. The baseline mission mode selected for MSR is Mars orbit rendezvous (MOR), analogous in concept to the lunar orbit rendezvous (LOR) mode used for Apollo in the 1960s. Specifically, MSR will employ two NASA-provided landers of nearly identical design and one CNES-provided orbiter carrying a NASA payload of rendezvous sensors, orbital capture mechanisms, and an Earth entry vehicle (EEV). The high-level concept is that the landers will launch surface samples into Mars orbit, and the orbiter will retrieve the samples in orbit and then carry them back to Earth. The first element to depart for Mars will be one of the two landers. Currently, it is proposed that an intermediate class launch vehicle, such as the Boeing Delta 3 or Lockheed Martin Atlas 3A, will launch this 1800-kg lander from Cape Canaveral during the May 2003 opportunity. The lander will utilize a Type-1 transfer trajectory with an arrival at Mars in mid-December 2003. Landing will be aided by precision approach navigation and a guided hypersonic entry to achieve a touchdown accuracy of 10 km or better. Although the exact landing site has not yet been determined, it is estimated that lander resource constraints will limit the site to between 15 degrees north and south latitudes. Following touchdown, the lander will deploy a six-wheeled, 60-kg rover carrying an extensive suite of instruments designed to aid in the analysis of the local terrain and collection of core samples from selected rocks. The surface mission is currently designed around a concept called the surface traverse. Each traverse will involve the rover exploring a selected area of terrain up to 100 meters from the lander, the collection of rock core samples, and the delivery of the samples from the traverse back to a sample canister on the lander. Planning estimates indicate that up to three traverses may be possible during the expected 90-sol lifetime of the lander. The canister that will receive the samples from the rover will be attached to the top stage of a small solid-fueled rocket mounted to the deck of the lander. This rocket is called the Mars Ascent Vehicle (MAV) and consists of three stages weighing a total of about 140 kg. After the conclusion of the surface mission, the MAV will lift-off and insert the sample canister into a near-circular orbit with an altitude of about 600 km and inclination of 45 degrees. The sample canister will wait in this orbit until it is retrieved by the orbiter sometime in early 2007. In August 2005, the second lander and a CNES-provided orbiter weighing 2700 kg will depart for Mars. Currently, it is proposed that a single Ariane 5 provided by CNES will launch both of these two elements onto a Type-2 transfer trajectory. Although the orbiter and lander will be launched together, they will separate shortly after injection and will fly to Mars as two independent spacecraft. However, both spacecraft will perform a maneuver between 10 and 15 days after launch so that their arrival times at Mars differ by between 12 and 24 hours. This scheme will reduce the operational complexity at the encounter date. A set of four 60-kg surface probes will ride piggyback on the orbiter to Mars. These CNES-provided probes are called Netlanders and will serve as surface stations for scientific investigations independent of the Mars Sample Return goals. Starting approximately one month prior to arrival at Mars, the orbiter will begin to release the Netlanders one at a time. Each release cycle will take several days, and will include time for precision navigation to execute one or two maneuvers that will target the Netlanders to their proper landing site. All four deployment cycles will be completed prior to 10 days before arrival. Both the orbiter and lander will arrive in late-July 2006. Upon arrival, the lander will perform a precision landing and surface mission similar in concept to the one that was executed during the 2003 opportunity. Although the landing site for the 2005 opportunity has not been selected, it is expected to be different from the 2003 site to enhance the diversity of the collected samples. The orbiter's arrival at Mars will be highlighted by the first use of aerocapture to insert a spacecraft into a capture orbit around another planet. The choice of aerocapture, as opposed to a propulsive orbit insertion, was considered mission enabling due to a reduction of over 2000 m/s in mission AV. Aerocapture will be targeted to produce a 250 km x 1400 km capture orbit with an inclination of 45 degrees. Current analysis indicates that achieving this goal will require approximately six minutes of flight deep in the atmosphere with a targeted periapsis of approach of about 43 km. After factoring into account the penalty for carrying a heat shield to survive aerocapture, the net savings compared to a propulsive orbital insertion amounts to several hundred kilograms.
We examine the evolution of Jupiter's orbit as the Sun evolves to form a planetary nebula. We use an orbital synchronization mechanism, which was proposed by Tassoul (1987, ApJ, 322, 856), and find that Jupiter is likely to deposit a substantial fraction of its orbital angular momentum, in spinning-up the Sun, when the latter is on the upper Asymptotic Giant Branch (AGB). This amount of angular momentum is likely to cause some small degree of axisymmetric mass loss from the Sun. In this case, the Sun will form an elliptical planetary nebula. The formation of elliptical planetary nebulae with the influence of massive planets and brown dwarfs, may explain the large fraction of elliptical planetary nebulae among the total number of planetary nebulae known today.
Both mechanical and fountain-effect pumps are being considered for use in the in-orbit resupply of superfluid helium to a number of scientific instrument systems. This paper presents a review of the operating characteristics of these pumps. Particular emphasis will be given to the different methods of evaluating the efficiency of these pumps and their effectiveness in a transfer system.