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This paper describes the architecture of the Data Interface Facility, (DIF), a key element in NASA's advanced communications systems, and its core element, the Return Link Processor (RLP). Current activities to prototype some of the primary functions of the RLP are presented. It is shown how the prototype elements can be integrated into a full-performance DIF RLP.
Meteoroid and space debris impact data obtained from space exposed spacecraft parts or meteoroid experiments is analyzed. The returned space exposed surfaces considered include spacecraft windows, meteoroid experiments exposed on space flights, and samples of opportunity such as foil wrapped around a cosmic ray package. The debris flux is found to exceed the meteoroid flux in creating impact craters smaller than 20 microns in diameter, and may again exceed it for impact structures larger than a few mm in diameter. For impact structures between 100 microns and 1 mm in diameter, the debris flux is found to be several times less than the meteoroid flux. Determining more accurately the meteoroid or orbital debris mass that makes a given crater or hole size remains is considered to be the main remaining problem.
This paper presents a preliminary cost estimate for the design and development of the Mars Rover/Sample Return (MRSR) mission. The estimate was generated using a modeling tool specifically built to provide useful cost estimates from design parameters of the type and fidelity usually available during early phases of mission design. The model approach and its application to MRSR are described.
A 0.02-scale model of a lifting-body concept for possible application to the Assured Crew Return Vehicle from Space Station Freedom was tested at Mach 6 and 10 in air. Thermal mappings and surface streamline patterns were obtained at angles of attack ranging from 0 to 30 deg and unit Reynolds numbers Re from 2 to 8 x 10 to the 6th/ft. Areas that experienced the highest heating were near the model nose and tip-fin leading edges. The effect of Re on windward centerline heating coefficients was negligible, whereas increases in angles of attack produced increases in heating. At Mach 6 and the highest unit Re, turbulent heat at the windward centerline was three to four times the laminar level. Leeward crossflow separation and vortex reattachment along the centerline are evident across the Re and angle-of-attack ranges tested, indicative of a complex flowfield.
The effects of Mach number, Reynolds number, and ratio of specific heats on the aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration were examined for a range of angles of attack from -5 deg to 50 deg. Predictions made with a Langley-developed, three-dimensional Navier Stokes solver known as LAURA, which was exercised as an Euler solver for the present study, are compared with the experimental results. Unlike the Shuttle Orbiter, which experienced a significant nose-up increment in pitching moment with decreasing specific heat ratio (i.e., real gas effects), the aerodynamic characteristics of this lifting-body configuration are insensitive to changes in specific heat ratio. The maximum trimmed lift-to-drag ratio achieved was about 1.5. Predicted inviscid values of aerodynamic coefficients were generally in good agreement with measurement.
The current reference Mars Rover and Sample Return mission is described. Technical issues are outlined, including high-resolution image acquisition and reconstruction, approach navigation, ground and flight systems operational complexity, rover autonomy, autonomous rendezvous and docking in Mars orbit, aerocapture and aeromaneuver, estimating the probability of mission success, and end-to-end information system design. Focus is placed on lander hazard identification and avoidance, pinpoint landing guidance and control, Mars ascent vehicle guidance and control, planetary protection and quarantine, sample acquisition and preservation, project management and control, systems requirements and interface control, and costing. In addition, program issues such as international participation, fiscal constraints, and launch-vehicle availability are considered.
ICE experienced a successful encounter with Comet Giacobini-Zinner on September 11, 1985. The comet-intercept trajectory, earth-return trajectory, and two cases of a possible 2014 capture trajectory are presented. Also, the essential propulsive trajectory correction maneuvers and the projected August 10, 2014 lunar swingby maneuver are discussed. The targeting methodology used for designing the trajectories is also discussed.
The principal objectives of the USO experiment on the COBE spacecraft are defined, and results of space qualification studies for the COBE USO experiment are summarized. The principal objectives of the experiment are: (1) to determine flight performance of the USO coupled to the second-generation TDRSS transponder; (2) space qualify TDRSS noncoherent one-way return-link Doppler tracking; and (3) analyze algorithms for one-way navigation with real data. The three objectives of the experiment have been met in the first stage of the experiment analysis.
Trajectory design issues beyond those under consideration in the Aeroassist Flight Experiment (AFE) program and more in line with the Shuttle design concept are presented. This is deemed necessary as flight regimes encountered in Lunar/Mars aerocapture trajectories at Mars and return-to-earth will provide more stringent requirements on vehicle design than AFE and are not exactly quantifiable at this time. Investigations are conducted of areas where explicit trajectory control is practical and can most directly influence constraints critical to vehicle design for manned and unmanned interplanetary missions. Primary consideration is given to the ability to control vehicle heating during entry, and this particular area of Shuttle entry design is discussed in detail.
The current space exploration initiative is focused around the development of the Space Station Freedom (SSF). Regular resupply missions must support a full crew on the station. The present mission capability of the shuttle is insufficient, making it necessary to find an alternative. One alternative is a reusable Cargo Return Vehicle (CRV). The suggested design is a biconic shaped, dry land recovery CRV with an advance recovery system (ARC). A liquid rocket booster will insert the CRV into a low Earth orbit. Three onboard liquid hydrogen/liquid oxygen engines are used to reach the orbit of the station. The CRV will dock to the station and cargo exchange will take place. Within the command and control zone (CCZ), the CRV will be controlled by a gaseous nitrogen reaction control system (RCS). The CRV will have the capability to exchange the payload with the Orbital Maneuvering Vehicle (OMV). The bent biconic shape will give the CRV sufficient crossrange to reach Edwards Air Force Base and several alternative sites. Near the landing site, a parafoil-shaped ARS is deployed. The CRV is designed to carry a payload of 40 klb, and has an unloaded weight of 35 klb.
NASA is committed to placing a permanent space station in Earth orbit in the 1990's. Space Station Freedom (SSF) will be located in a 220 n.m. orbit at 28.5 degrees inclination. The Winged Cargo Return Vehicle's (CRV) primary mission is to support SSF crew by flying regular resupply missions. The winged CRV is designed to be reusable, dry land recoverable, and unmanned. The CRV will be launched inline on three liquid hydrogen/oxygen rocket boosters with a payload capacity of 113,000 lbs. The three boosters will take the CRV to an orbit of 50 by 110 n.m. From this altitude the orbital manuevering engine will place the vehicle in synchronous orbit with the space station. The winged CRV will deliver cargo modules to the space station by direct docking or by remaining outside the SSF command zone and using the Orbital Maneuvering Vehicle (OMV) to transfer cargo. After unloading/loading, the CRV will deorbit and fly back to Kennedy Space Center. The CRV has a wing span of 57.8 feet, a length of 76.0 feet, and a dry weight of 61.5 klb. The cargo capacity of the vehicle is 44.4 klb. The vehicle has a lift-drag ratio of 1.28 (hypersonic) and 6.0 (subsonic), resulting in a 1351 n.m. cross range. The overall mission length ranges between 18.8 and 80.5 hr. The operational period will be the years 2000 to 2020.
Described herein is a series of design studies concerning the Assured Crew Return Vehicle (ACRV). Study topics include a braking and landing system for the ACRV, ACRV growth options, and the design impacts of ACRV's role as a medical emergency vehicle. Four alternate designs are presented for the ACRV braking and landing system. Options presented include ballistic and lifting body reentries; the use of high-lift, high-payload aerodynamic decelerators, as well as conventional parachutes; landing systems designed for water landings, land landings or both; and an aerial recovery system. Uses of the ACRV or a similarly designed vehicle in several roles for possible future space missions are discussed, along with the required changes to the ACRV to allow it to perform these missions optimally. The impacts on the design of the ACRV due to its role as an emergency vehicle were studied and are presented here. This study included the design of a stretcher-like system to transport an ill or injured crewmember safely within the ACRV; a compilation of necessary medical equipment and decisions on how or where to store it; and recommendations about internal and external vehicle characteristics that will ease the transport of the ill or injured crewmember and allow for swift and easy ingress/egress of the vehicle.
Thermochemical relaxation distances behind the strong normal shock waves associated with vehicles that enter the Earth atmosphere upon returning from a manned lunar or Mars mission are estimated. The relaxation distances for a Mars entry are estimated as well, in order to highlight the extent of the relaxation phenomena early in currently envisioned space exploration studies. The thermochemical relaxation length for the Aeroassist Flight Experiment is also considered. These estimates provide an indication as to whether finite relaxation needs to be considered in subsequent detailed analyses. For the Mars entry, relaxation phenomena that are fully coupled to the flow field equations are used. The relaxation-distance estimates can be scaled to flight conditions other than those discussed.
Lightning properties obtained by a network of magnetic direction finders and by electric field measurements for distances from 50 to 500 km are compared for three summer thunderstorms in Sweden. The data from direct field recordings indicate 31, 17, and 26 pcts. of negative subsequent return strokes with peak current (as inferred from the peak electric field) higher than the first. Electric fields from first strokes are compared with normalized amplitudes registered by the magnetic direction finding system. The efficiency of detection by the magnetic direction finding system is discussed in terms of the percentage of lightning flashes observed by electric field measurements that are not localized. Statistics of the number of strokes per flash and the interstroke time intervals are presented.
The design of a Mars Rover Sample Return (MRSR) mission that satisfies scientific and human exploration precursor needs is described. Elements included in the design include an imaging rover that finds and certifies safe landing sites and maps rover traverse routes, a rover that operates the surface with an associated lander for delivery, and a Mars communications orbiter that allows full-time contact with surface elements. A graph of MRSR candidate launch vehice performances is presented.
This paper explains why returned samples will be critical elements of future Mars exploration. The types of samples that are needed and how they will be studied is addressed. Why putative Martian meteorites are not acceptable substitutes for documented samples is pointed out.
It is argued that returned comet samples should be distributed to individual laboratories for analysis much in the same way that lunar samples are. In the context of emerging technologies, certain developments in other fields which may prove beneficial to the study of comets are reviewed, including cryogenic and tomographic techniques. It is noted that the lack of expertise in the area of cryogenic examination of comet samples is a current weakness in the ability of the extraterrestrial material community to handle comet samples and thus requires further training and preparations. Computer tomography scan images of a 'dirty snowball' are presented to illustrate the importance of developing new methods for comet sample analysis. It is suggested that microanalytic measurements of extraterrestrial samples at the atom-counting limit appear promising for the future.
The autonomous lunar precursor mission presently proposed for low cost manned revisits to the lunar surface, designated 'Lunar Oasis by Automation and Robotics', employs two self-similar robotic vehicles, and three experimental 'biovolumes' which are directly derived from the lunar landers that are initially entrusted with fulfilling mission hardware-delivery requirements. The two vehicles would support the experimental biovolumes by returning resources for use in physicochemical or biological processing. Delivered resources would be combined with those in situ to furnish reconfigured biovolumes with the capabilities for autonomous biological growth and wastes-reprocessing demonstrations.