Mission Payload Research Articles

Error analyses of CHAMP data for recovery of the Earth?s gravity field

A preliminary commission error analysis whereby orbit perturbation theory and other techniques are used to assess and predict the recovery of the Earth’s gravity field from the challenging microsatellite payload (CHAMP) mission is developed and implemented. With CHAMP launched in July 2000, accumulated evidence is now available to quantify the errors in the recovery procedure including the orbital precision from GPS, attitude errors, accelerometer noise and thruster mismatch/misalignment. For the latter, numerical integrations using a variable length single-step Runge–Kutta integrator and a fixed length multi-step method are compared to assess the error associated with assuming that the thruster misalignment can be spread uniformly across a step interval. Error degree variances from simulated studies are compared to results from a recently released CHAMP-based gravity field, EIGEN-1S. It is seen that the orbital positioning, as derived from the onboard GPS receiver, is critical, with accelerometer noise contributing at a lower level. Attitude error, at currently quoted accuracy, is not significant as an error source.

Melting processes under microgravity conditions

The kinetics of melting pivalic acid (PVA) dendrites was observed under convection-free conditions on STS-87 as part of the United States Microgravity Payload Mission (USMP-4) flown on Columbia in 1997. Analysis of video data show that PVA dendrites melt without relative motion with respect to the quiescent melt phase. Dendritic fragments display shrinking to extinction, with fragmentation occurring at higher initial supercoblings. Individual fragments follow a characteristic time-dependence derived elsewhere. The microgravity melting kinetics against which the experimental observations are compared is based on conduction-limited quasi-static melting under shape-preserving conditions. Agreement between analytic theory and our experiments is found when the melting process occurs under shape-preserving conditions as measured using the C A ratio of individual needle-like crystal fragments.

Experiment configuration and preliminary results of two-phase fluid loop experiment (tpflex)–sts-85 mission payload

A two-phase fluid loop (TPFL), which enables removal of a large amount of heat and long-distance heat transportation utilizing latent heat of the pumped working fluid, is expected to be an essential thermal control system for the future spacecraft and orbital infrastructure. Two-Phase Fluid Loop experiment (TPFLEX) on Space Shuttle (STS-85) was conducted in order to demonstrate the operational principle, DP(=vapor pressure in the evaporator minus liquid pressure in the evaporator)>0, of TPFL under micro-gravity. As a result of this experiment we obtained the DP data as a function of heat load on the evaporator for a fixed flow rate.

7.7.4 A System Approach to the Conceptual Design of an Upgrade Mission Payload for Maritime Helicopters

AbstractThe global budget climate has laid restrictions on the design and development of new military aircraft to meet the demands of additional mission capabilities. To enhance mission capabilities, mid‐life upgrade of in‐service aircraft, involving placement of advanced technology mission systems onboard, is acknowledged as a cost‐effective option. To facilitate the mid‐life upgrade process, a “Decision Support System” is required to identify the state‐of‐the‐art mission systems that will provide the required mission capabilities. An “Integrated Decision Support System” (IDSS) that analyses the mission requirements holistically, to identify mission systems that enhance the capability of the aircraft is presented in this paper. A brief outline of the IDSS framework developed by a systems approach is presented initially, followed by a detailed discussion of the ‘Mission Payload Design’ (MPD) sub‐model. The MPD sub‐model prioritises the identified mission systems based on their relative dependence and degree of contribution in mission accomplishment, to formulate mission payload alternatives for aircraft upgrade.

High rate low latency GPS ground tracking network for CHAMP

High Rate / Low Latency GPS ground measurements for the GPS occultation experiment of the Challenging Minisatellite Payload (CHAMP) mission are provided by a dedicated global tracking network. The network is being established jointly by GFZ and Jet Propulsion Laboratory (JPL) to ensure adequate ground data availability for processing of globally distributed GPS occultation events. GFZ operates a sub-network, which is currently composed of 11 stations. This paper discusses various scientific and technical issues such as network topology design, requirements for an autonomous and continuously operating GPS ground tracking station, near real time data transfer, interfaces and operational data flow.

Precision Spacecraft Attitude Estimators Using an Optical Payload Pointing System

The advancement of satellite program technology mandates that issues affecting future spacecraft system development cost be resolved, while continuing to meet smaller size, lower power, and more stringent mission requirements. In the specie c area of the spacecraft attitude determination system, an optical payload pointing system can servea dual purpose: to support the payload mission and to providepreciseattitude information for the spacecraft functions. As a result, star trackers orotherprecisereferencesystems can beeliminated to reduce the development cost. Additional benee ts include performance robustness to spacecraft motion or disturbance and potential cost saving due to weight reduction, e.g., launch cost. Attitude estimators using an optical payload pointing system and strap-down gyros are derived in detail for two distinct formulations: spacecraft body and inertial formulations. Alternative Kalman e lter equation derivations are shown using the skew symmetric matrix properties of the attitude kinematics equation. It is also shown that both formulations produce identical solutions. However, the body formulation is conceptually easier to understand, and the inertial formulation requires less processing time. To show thespacecraft attitudeestimatorperformanceusing an optical payload pointing system, an exampleisshown for an agile low-Earth-orbit satellite.

Technical evolutions of the french multipurpose instruments for cognitive neurosciences

Since the first French flight in space in 1982, the CNES has developed a wide range of instruments, especially in the field of Neurosciences. The design of these instruments has considerably evolved from rather simple equipment up to much more sophisticated tools that are being specially tayloried for these missions. Four major phases can be identified: • - a simple adaptation of an echographe leading to the first neurosciences experiments (the ARAGATZ'88 mission), • - the ILLUSIONS and VIMINAL instruments used during the ANTARES'92 and ALTAIR'93 missions, • - the COGNILAB instrument developed for the CASSIOPEE'96 mission, to be re-used in 1997 and in 1999, • - a preliminary design of the 1999 mission payload, including virtual reality concepts, in a modular design to adapt to the European COF. Aside from the evolution of scientific requirements, the experience gained during the flights led to progressive improvements in the different technical parts, including visual system, body restraint systems, accessories, such as a force feedback joystick, computer and software, etc. This paper describes the technical evolutions in the CNES Neurosciences program.

Effect of residual accelerations during microgravity directional solidification of mercury cadmium telluride on the USMP-2 mission

Directional solidification of mercury cadmium telluride (MCT) requires that the temperature gradient to growth rate ratio be high to avoid constitutional supercooling. With the optimum gradient condition for solidifying MCT in NASA's advanced automated directional solidification furnace (AADSF), it is necessary to use translation rates as low as 0.2 μm/s. The result is that any fluid flow with a velocity comparable to or higher than this will dominate the solidification characteristics, particularly the compositional distribution in an alloy such as this which has a large solidus-liquidus separation. In an effort to reduce fluid flow velocities a space experiment was performed. On the second United States Microgravity Payload Mission (USMP-2), the AADSF made its maiden flight and successfully completed growth of a MCT boule 16 cm long. The furnace was located approximately 3 m away from the center of gravity of the space shuttle, and this combined with the drag component of residual acceleration present during flight, resulted in quasisteady residual accelerations of the order of 1 μg 0 where g 0 is the earth's natural gravity. Of more importance is that different orbiter attitudes during the mission produced significant differences in the resultant residual acceleration vector, in both magnitude and direction and that these differences caused large compositional variations both across the radii of the boule and along the surfaces of the boule. Comparison will be made with examples grown on the ground and in magnetic fields.

The AMICIST auroral sounding rocket: A comparison of transverse ion acceleration mechanisms

Recent auroral sounding rocket data illustrate the relative significance of two transverse acceleration of ion (TAI) mechanisms for initiating nightside auroral ion outflow. First, the new data from this two payload mission show clearly that for lower hybrid solitary wave events, a) these individual events are spatially localized to scales approximately 100 m wide perpendicular to B, and b) the probability of occurrence of the events is greatest at times of maximum VLF wave intensity. Second, ion acceleration by broadband, low frequency electrostatic waves is observed in a 30 km wide region at the poleward edge of the arc. The fluxes from this and other sounding rockets are shown to be consistent with DE‐1 and Freja outflow measurements, indicating that the AMICIST observations show the low altitude, microphysical signatures of nightside auroral outflow from TAI.

Fast Attack Combatant for Near‐Shore Operation

ABSTRACT The technologies of advanced naval engineering were used to design a corvette‐size vessel which transited the Atlantic Ocean in 1992, without refueling, at an average speed of 53.1 knots. The application of these and related technologies to fast attack craft (FAC) of 200 to 500 tonnes can result in vessels with extraordinary capabilities. The combination of 40 + knot speeds, navigational draft of less than 2 meters, enhancements for signature reduction, active and passive sensors, data linking with remote sources and 75 + tonnes for mission payload and fuel offer creative options for operation in a littoral environment.This paper describes a multi‐mission, 250‐tonne FAC, propelled by waterjets with power supplied by gas turbines and diesels. Sufficient detail is provided to indicate weight allocated for mission equipments, the weights for light ship and the loads and optional stations for launching/recovering delivery vehicles. Operational characteristics of this FAC are described such that they may be used for studying tactical advantages.

Isothermal dendritic growth experiment flight unit tests

The Isothermal Dendritic Growth Experiment is a microgravity materials science experiment planned to fly on three United States Microgravity Payload missions. Test of the engineering unit began in 1990. In mission simulations, succinonitrile dendrites were repeatedly solidified and photographed at supercoolings from 0.1 to 1.0 centigrade degrees. Then tip radii and velocities were measured to evaluate performance. Flight unit testing began in 1992. Flight unit teleoperational capabilities include uplink of commands and downlink of scientific and engineering data such as slow scan television images of solidifying dendrites and acceleration data. The first flight will be on the USMP-2 carrier in the Shuttle cargo bay during 1994.

Land mobile satellite communication system using the geostationary platform

If the geostationary platform (GPF) is available in the first decade of the 21st century, the demand on a voice communication base is expected to be about 800,000 Land Mobile Satellite Service (LMSS) terminals. To improve the frequency use of an LMSS over Japan, it is important to reuse the frequencies allocated to the Tokyo metropolitan area in the Osaka-Kobe-Kyoto region. An onboard multibeam antenna with beam intervals of 0.468 deg (L band 30-m class) for each of the 11 beams covering the main Japanese islands, cluster power feeding, and dual polarization can enable the reuse of frequencies at one-beam intervals. A deployable antenna is under development that is lightweight and has a high storage efficiency for launching. This will form the basis for an antenna with an aperture diameter of 30 m. Such a large antenna will be a key factor in reducing ground terminal size. This paper estimates the weight and power requirements of the mission payloads onboard the GPF, which include an LMSS, a millimeter wave personal satellite communication system, and a high-definition television/regional television satellite broadcasting service system. The total weight of this GPF would be approximately 4-5 tons.

The use of artificial intelligence techniques to improve the multiple payload integration process

A maximum return of science and products with a minimum expenditure of time and resources is a major goal of mission payload integration. A critical component then, in successful mission payload integration is the acquisition and analysis of experiment requirements from the Principal Investigator (PI) and Payload Element Developer (PED) teams. This article describes one effort to use Artificial Intelligence (AI) techniques to improve the acquisition and analysis of experiment requirements within the Payload Integration Process.

U.S. space station freedom: Orbital assembly and early mission opportunities

This paper describes the U.S. Space Station launch and assembly sequence and the mission support capabilities available in man-tended and permanently manned modes. Design and operational constraints affecting launch package configurations and orbital deployment and assembly are defined. The resulting flight configurations and associated payload accommodations are described with emphasis on the early flights through man-tended capability. The initial mission objectives will emphasize engineering data collection (structural, thermal, power, control, propulsion, etc.) to verify design and to calibrate system models and simulation tools. On-board information will be used to characterize the microgravity, electromagnetic, particulate, and other environmental features for the benefit of Space Station users. Early flight configurations will provide accommodation for attached payloads which may include experimentation in materials exposure, solar physics, and Earth and stellar observations. These mission payloads, once installed and initialized by the Shuttle flight crews, will be operated by ground control from their mission specialists. When the pressurized laboratory is added, the mission emphasis will shift to microgravity experimentation including material processing and manufacturing research, life sciences, and supporting technology research and development. These missions will be operated in a man-tended mode. Installation, real-time operation, and servicing will be performed by Shuttle during periodic visits; a combination of ground control and automatic modes will be used between Shuttle flights. As Station build-up continues, resources available for payloads will also increase. With the achievement of the permanently manned capability, continuous crew support will be available for mission activities. This paper summarizes each step of this build-up process and the capabilities available for Station operation and payload activities versus time.

On-orbit experiment plan on solar thermodynamic power system

This paper presents the on-orbit experiment plan of the solar thermodynamic power system operating the space flyer unit (SFU) presently being developed in Japan. The on-orbit experiments will be conducted with four primary objectives. First, to demonstrate the operation of the solar thermodynamic system in space environment. Second, to assess the performance capability of the solar thermodynamic system in space. Third, to verify functions of thermodynamic engine, solar ray collector, thermal storage system and space radiator especially from view point of durability and reliability. Fourth, to examine how vibration and inertia caused by moving parts of thermodynamic engine affect the primary functions of SFU and other mission payloads. The solar thermodynamic power system consists of a free piston-type Stirling engine, a linear synchronous generator, a Cassegrain-type solar ray collector, a thermal storage system and a space radiator. This system produces the continuous electric power of 2.5 kW, 330 V a.c. using 4.3 m dia solar collector. The output power is fed to the MPD thruster experiment. This experiment is proposed for the second SFU mission called “Energy Mission” in the mid 1990s. As a preliminary experiment, the ballistic flight experiment using sounding rocket is also planned.

Efficient sizing of a cargo rotorcraft

A new cargo rotorcraft sized to meet the most demanding of the future combat airlift requirements being postulated by the U. S. Army could be unnecessarily large and expensive. Overall requirements could be satisfied more efficiently with a smaller aircraft by using multilift for infrequent peak demands. This paper describes a methodology for assessing the benefit of multilift and the influence of usage spectrum on required aircraft size and achievable overall productivity. For characteristic distributions of mission payload, radius of action, and ambients, it is shown that aircraft design gross weight could be reduced as much as 40% and overall productivity improved as much as 50% by using multilift for less than 5% of the missions.

Pulsed plasma thrusters for orbit transfer

Introduction T HE use of electric propulsion for near earth missions has recently received increased attention with the recognition of possible long-duration missions involving payloads that require substantial levels of electrical power. If most of the mass of the power supply (including heat rejection) can be ascribed to the mission payload, then the optimum exhaust velocity can increase substantially above levels accessible with chemical thrusters. From classic mission analyses for electric propulsion, the optimum exhaust speed is given by u= (kectT)' where k=l-2 depending on the ratio of u to the necessary velocity increment At>; e is the efficiency of thrust power from total power, a. is the specific power associated only with the electric propulsion system, and r is the mission time. If the total power system has a specific power of 30 W/kg for example, but 80% of this system is needed by the mission, then a =150 W/kg. A thirty-day mission (with k = 2 and e = 0.5) would then have an optimum specific impulse of 2000 s. There are several electric propulsion concepts that could satisfy this specific impulse requirement. This Note discusses some aspects of a particular approach, based on a system that has already performed successfully (albeit modestly) on long term space missions, the pulsed plasma microthruster (PPT). The principal interest is the introduction of electrical propulsion in a manner that can evolve as electrical power levels increase, drawing on actual flight experience while maintaining the physical processes in the thruster. Such introduction would benefit from experience with the PPT design. That is, optimization of thruster performance can be sacrificed to system simplicity in order to match available power supplies and minimize risk. For example, the PPT uses a Teflon fuel block fed into the electrode region by a negator spring, and thus avoids repetitive valve operation and cryogenic fuel storage. Furthermore, arc discharges tend to operate at voltages that are independent of the charging voltage of the current source. The energy from a pulsed plasma thruster might vary slightly from shot to shot, but the requirement for voltage regulation of the power supply is much less severe than for other types of electric thruster. This Note reviews some aspects of mission performance using pulsed plasma thrusters. Recent experimental work suggests the possibility of a single thruster that could match increasing space-electric power levels.

The german microgravity program: Objectives and present status

Within the space program of the Federal Republic of Germany the microgravity program in connection with the utilization of SPACELAB constitutes a central task which determines the long-term program concepts and also their relation to German participation in future ESA programs. The scientific preparatory programs under way for some years now have made further progress. Extensive flight experience and valuable scientific results were obtained on the basis of successful rocket pre-programs. The present paper describes the process in which scientific and organisational priorities are being defined for the planning and execution of the experimental programs. In order to obtain a sufficient number of flight opportunities, payloads for SPACE SHUTTLE missions, in particular under the NASA GAS Program, as well as experimental equipment such as the materials laboratory (MSDR) for FSLP are being developed. The German program focuses on preparing a German SPACELAB mission D1 planned for 1985, which is intended to verify the applicability and efficiency of manned research laboratories for industry and the scientific community. A second emphasis is on preparing the use of SHUTTLE-supported re-usable space platforms.

CTOL/VSTOL Comparison-A View from the Deck

Aircraft carrier capabilities are compared for catapult launched and arrested landing aircraft—CTOL aircraft, and short takeoff and vertical landing aircraft—V/STOL aircraft. The V/STOL aircraft are capable of takeoff and landing in the vertical mode with increased mission payload and performance when operated in an STO (short takeoff) overload condition. Investigations were conducted on how future generic aircraft were affected by the ship environment and how aircraft operations could be optimized. Starting with the CTOL aircraft, investigations were performed to determine the limiting operational factors and constraints. This established baseline parameters for the operational sensitivity studies. Sensitivity studies were then made comparing equal size and equal cost airwings. Results of these studies parametrically show the impact on sorties over all mission ranges out to 700 n.mi. at various cost levels.

Exosat for Ground-based Observers

The first version of this mission was approved by the European Space Agency (ESA) Council in 1973 but did not infact start its Phase B study until 1977. The early baseline design had been constrained by the requirement to make the spacecraft compatible in size and weight with the performance of a NASA Delta rocket, since this was seen as a necessary back-up to Ariane, then at an early stage of development. The mission payload then evolved with time due to the changing role dictated by the technical successes and observations made by the series of well-known X-ray satellites. The final spacecraft has emerged to have a unique capability since all the other X-ray satellites except the small Hakucho have now expired.