Geometric aspects of ground station/satellite communications
Geometry aspects of communications between ground station and satellite in elliptical orbit
GEOMETRY↗
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Dynamics of momentum biased spacecraft in a near-polar orbit
The equations of motion of a momentum biased spacecraft are derived in a general form. The spacecraft is assumed to be orbiting in a near-elliptical orbit. An aerodynamic torque model which accounts for the atmospheric superrotation is assumed. The equilibrum attitude angles are obtained in terms of modified Bessel functions. Analytic expressions for the long-term motion of the momentum biased axis are derived for special cases. The analysis is applicable to the Magsat mission.
Sellappan, R. G.↗
Differential GPS approaches to orbit determination of high-altitude earth satellites
Orbit determination of earth satellites using conventional (up-looking) differential GPS is limited to lower altitudes due to the finite antenna beamwidth and altitude of GPS satellites. Higher user satellites suffer from having short GPS view period; at sufficiently high altitudes, up-looking GPS visibility becomes so poor that no differential measurements are possible. For such higher user satellites, two alternative differential GPS approaches are suggested. A covariance analysis is performed investigating the capabilities of these approaches at various user satellite altitudes. Best differential GPS data types are identified for satellites within different altitude ranges. Error sources and other factors affecting the orbit determination accuracy are examined. Relative merits of these approaches are compared. When applied to QUASAT, the proposed earth orbiting radio telescope with an elliptical orbit, a sub-meter orbit determination accuracy is expected.
Wu, S.-C.↗
Analytical Evaluation of a Method of Midcourse Guidance for Rendezvous with Earth Satellites
A digital-computer simulation was made of the midcourse or ascent phase of a rendezvous between a ferry vehicle and a space station. The simulation involved a closed-loop guidance system in which both the relative position and relative velocity between ferry and station are measured (by simulated radar) and the relative-velocity corrections required to null the miss distance are computed and applied. The results are used to study the effectiveness of a particular set of guidance equations and to study the effects of errors in the launch conditions and errors in the navigation data. A number of trajectories were investigated over a variety of initial conditions for cases in which the space station was in a circular orbit and also in an elliptic orbit. Trajectories are described in terms of a rotating coordinate system fixed in the station. As a result of this study the following conclusions are drawn. Successful rendezvous can be achieved even with launch conditions which are substantially less accurate than those obtained with present-day techniques. The average total-velocity correction required during the midcourse phase is directly proportional to the radar accuracy but the miss distance is not. Errors in the time of booster burnout or in the position of the ferry at booster burnout are less important than errors in the ferry velocity at booster burnout. The use of dead bands to account for errors in the navigational (radar) equipment appears to depend upon a compromise between the magnitude of the velocity corrections to be made and the allowable miss distance at the termination of the midcourse phase of the rendezvous. When approximate guidance equations are used, there are limits on their accuracy which are dependent on the angular distance about the earth to the expected point of rendezvous.
Eggleston, John M.↗
A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer
In the present day, orbit determination by Global Positioning System (GPS) is not unusual. Especially for low-cost small satellites, position determination by an on-board GPS receiver provides a cheap, reliable and precise method. However, the original purpose of GPS is for ground users, so the transmissions from all of the GPS satellites are directed toward the Earth s surface. Hence there are some restrictions for users above the GPS constellation to detect those signals. On the other hand, a desire for precise orbit determination for users in orbits higher than GPS constellation exists. For example, the next Japanese Very Long Baseline Interferometry (VLBI) mission "ASTRO-G" is trying to determine its orbit in an accuracy of a few centimeters at apogee. The use of GPS is essential for such ultra accurate orbit determination. This study aims to construct a method for precise orbit determination for such high orbit users, especially in High Elliptical Orbits (HEOs). There are several approaches for this objective. In this study, a hybrid method with GPS and an accelerometer is chosen. Basically, while the position cannot be determined by an on-board GPS receiver or other Range and Range Rate (RARR) method, all we can do to estimate the user satellite s position is to propagate the orbit along with the force model, which is not perfectly correct. However if it has an accelerometer (ACC), the coefficients of the air drag and the solar radiation pressure applied to the user satellite can be updated and then the propagation along with the "updated" force model can improve the fitting accuracy of the user satellite s orbit. In this study, it is assumed to use an accelerometer available in the present market. The effects by a bias error of an accelerometer will also be discussed in this paper.
Ikenaga, Toshinori↗
Aerocapture as an Enabling Technology for Planetary Missions
Aerocapture is an atmospheric maneuver where aerodynamic forces are used to transfer a spacecraft from a hypersonic orbit to a targeted capture orbit. Aerocapture provides large mass benefits over all-propulsive maneuvers that are typically used to enter capture orbits, as the aerodynamic forces of the vehicle in an atmosphere slow down the vehicle rather than using propellant. Additionally, aerocapture is a quick maneuver where the spacecraft enters the target orbit after a short transit through the atmosphere versus similar aero-assist maneuvers like aerobraking, where the vehicle enters the desired orbit incrementally from a highly elliptical orbit and usually takes a time period of the order of months.
Soumyo Dutta↗
Optimal switching in coplanar orbit transfer
Weierstrass condition for optimality in coplanar elliptic orbit transfer involving velocity impulse control variables, discussing switching relations at corner of optimal trajectory
Culp, R. D.↗
Optimum deorbit positioning for minimum impulse reentry.
Optimum deorbit positioning on planetocentric elliptic orbit for single impulse reentry treated by two coupled quartic polynomials
ELLIPTICAL ORBIT↗
The solar electric propulsion stage concept for high energy missions.
Definition of multimission and engine performance requirements for candidate solar electric propulsion stage configurations, considering launch vehicle compatibility, electric propulsion integration, payload requirements, and the effects of environmental extremes. Electric propulsion power options include two solar array power levels (15/22 kW), up to twelve electric thrustors of 30 cm diameter and 2.7 kW each, five to eight power conditioning units, and a maximum mercury propellant capacity of 1530 kg. In performance, the stage with a dry weight of 700 to 900 kg can deliver a net mass of 756 kg into Saturn orbit, 329 kg into a tight Mercury orbit, and 334 kg within 0.1 AU of the sun. The stage can also deliver a round trip payload of 3350 kg to geosynchronous orbit and return from an intermediate elliptical orbit using the Shuttle/Tug. Thus, a versatile stage is developed which competes effectively in performance with existing integrated spacecraft and promises considerable savings in total program costs.
Guttman, C. H.↗
Orbit crowding of molecular gas at a bar-spiral arm transition zone in M83
The southwestern bar-spiral arm transition zone in M83 is been studied in CO, H-alpha, H I, red light, and the radio continuum. A massive molecular gas complex in the heart of the transition zone is composed or two principal components which have the morphology and kinematics expected from orbit crowding, where gas on highly elliptical orbits form the bar region converges with gas on more circular orbits from the spiral arm region. Three mechanisms for the origin of the orbit crowding are investigated, and it is proposed that the crowding is due primarily to density wave streaming motions caused by the bar and spiral arms. The inner CO component is partially coincident with a region of highly polarized radio continuum emission which precedes the H-alpha spiral arm by 15-25 arcsec, indicating that it lies on or just downstream from a shock front. This suggests that the bar gas approaching the transition zone is shocked and explains the ridge of dense gas seen upstream from the spiral arm.
Kenney, Jeffrey D. P.↗
Lunar Relay Coverage Analysis for RF and Optical Links
This paper describes our effort in searching different lunar relay architectures and assessing their coverage performance and other pros and cons for RF and optical links. Regarding RF links, we considered three notional lunar relay architectures, all communicating with the three sites of the Deep Space Network (DSN): Goldstone, California; Canberra, Australia; and Madrid, Spain: 1) A constellation of three relay orbiters: two at the Polar frozen elliptical orbits, and one at an equatorial circular orbit. 2) One relay orbiter in a 74-day Lissajous orbit at Earth-Moon Lagrange Point L2. 3) One relay orbiter in a 14-day Lunar distance retrograde orbit. Regarding optical link, we considered a notional multi-hop relay architecture that consists of one relay orbiter in a 74-day Lissajous orbit at Earth-Moon L2 (as in architecture #2 in the RF case), three Near-Earth relay satellites in geosynchronous orbit at longitudes 187° E, 110° E, and 347° E, and three ground stations at White Sands (US), Guam (US), and Tenerife (Spain). We introduce the additional constraints of Sun-“Earth”Probe (SEP) angle for return links, and Sun-Probe-“Earth” (SPE) angle for forward links. SPE and SPE angles determine the amount of sunlight (noise) that goes into the detector of the telescope, thus affecting the capacity of an optical link.
Lee, Charles↗
Sa-8 operational trajectory
Trajectory optimization for Saturn I /SA-8/ LAUNCH vehicle to place S- IV stage and payload in elliptical orbit
SATURN I /SA-8/ LAUNCH VEHICLE↗
Nonlinear studies, optimal control
Attitude control for earth satellite in elliptic orbit, linearization in attitude control, and computation of optimal controls by method based on second variations
LINEARIZATION↗
A Study of the Positions and Velocities of a Space Station and a Ferry Vehicle During Rendezvous and Return
A study is made of the families of non-thrusting ascent trajectories of a ferry vehicle during rendezvous with an orbiting body, referred to as a space station. It is shown that these trajectories may also be interpreted as descent trajectories of the ferry from the station to the earth. The rendezvous trajectories start at the end of the boost period (assumed to be 60 miles) and terminate at the station. The equations of motion are derived and results are shown for two typical orbits of the station: a 300-mile circular orbit and a 100-to-500-mile elliptical orbit. Trajectories are described in terms of a rotating coordinate system fixed in the station and launch conditions are tabulated in terms of non-rotating inertial coordinates. Boundaries are given in terms of launch (at time of booster burnout) and rendezvous conditions for the example cases. The considerations used to calculate these boundaries and the significance of some of the trends are discussed.
TRAJECTORY↗
A simplified explicit guidance scheme for orbital injection.
Onboard computer equations for rocket vehicle guidance to elliptical orbit, discussing iteration absence from computations and digital simulation results
Scofield, H. N.↗
Requirements and capabilities for planetary missions. Venus orbiter imaging radar 1983, volume 3
Two spacecraft will be launched in mid-1983 and inserted into a circular polar orbit around Venus about 6 months later. Elliptical orbits are also under consideration. The objective of the mission is imagery of the planet, at about the 200-m resolution level, with continuous altimetry and topographical studies. Science investigations will determine surface characteristics of the planet, study the surface/atmosphere interactions, and determine Venus' mass distribution. A plausible vehicle is based on a Mariner Jupiter/Saturn derivative; others being considered include a Lunar Polar Orbiter derivative and a Pioneer Venus Orbiter derivative.
Kindt, D. H.↗
Effects of solar radiation on the orbits of small particles
A modification of the Robertson (1937) equations of particle motion in the presence of solar radiation is developed which allows for partial reflection of sunlight as a result of rapid and varying particle rotations caused by interaction with the solar wind. The coefficients and forces in earlier forms of the equations are compared with those in the present equations, and secular rates of change of particle orbital elements are determined. Orbital dimensions are calculated in terms of time, probable sizes and densities of meteoric and cometary particles are estimated, and times of infall to the sun are computed for a particle moving in an almost circular orbit and a particle moving in an elliptical orbit of high eccentricity. Changes in orbital elements are also determined for particles from a long-period sun-grazing comet. The results show that the time of infall to the sun from a highly eccentric orbit is substantially shorter than from a circular orbit with a radius equal to the mean distance in the eccentric orbit. The possibility is considered that the free orbital kinetic energy of particles drawn into the sun may be the energy source for the solar corona.
Lyttleton, R. A.↗
In orbit manufacture of solar reflector satellites
The implementation of an idea reported by Oberth (1929) is discussed. This idea is concerned with the reflection of sunlight to the ground by means of orbiting metal mirrors. The approach used for implementing Oberth's idea makes use of aluminized Mylar as a reflecting material which is supported with the aid of a structure which includes six booms and a peripheral mirror cable. Four low resistance one turn coils are formed by the structure. Magnetic forces for attitude maneuvers are obtained by energizing the coils. A bipropellant propulsion system is to be used to raise the initial 330 km circular parking orbit to a 330 by 1000 km elliptic orbit. Two thrusters are employed. They are oriented parallel to each other and the mirror surface. The mirror satellite can be completely folded so as to maximize the launch capability of the Shuttle.
Muller, R. M.↗