High Earth Orbit Research Articles

Celestial Navigation Methods for Space Explorers

With the development of the space exploration missions, a practical and reliable capability for autonomous navigation is urgently needed to reduce operation cost, to improve operational efficiency and to increase mission safety by avoiding dangerous delay in ground control. Celestial navigation is a very attractive autonomous navigation solution for the space exploration missions. It suits LEO satellites as well as high earth orbit (HEO) satellites and deep space explorers. In this paper, a comprehensive overview of the current state and future trend of celestial navigation methods for three types of space explorers are introduced, which include satellites, deep space probes and planetary rovers.

Ultra long orbital tethers behave highly non-Keplerian and unstable

Large twin tethers are investigated as possible competitive-cost tools for non-gasdynamic descent, landing, takeoff and return from target celestial bodies and as passive tools for debris retrieval from orbit. The particular behavior of orbiting bodies connected with long cables is a recent preoccupation in astrodynamics and proves being full of unexpected results. The investigation here presented is focused on the non-Keplerian behavior of such large tether systems, considered in a first approximation as rigid or very stiff and massless. The investigation starts with the feasibility of non-gasdynamic orbital deployment of twin tethers without any involvement of expensive rocket propulsion means. The free tether release systems are associated to a horizontal impulsive separation (HIS) and eventual friction-free deployment to the desired length. This horizontal deployment seems to supply the most productive means of continuous separation and departure of masses in orbit. The relative motion during separation is studied and the observation is made that a considerable kinetic moment of the system preserves during all eventual phases of the flight. After the friction-free deployment the extending cable is instantly immobilized at the so-called connection moment. From here after the tether length remains constant. The evolution of the deployed tether is followed in order to record the specific behavior when the length of the tether is extremely great. The motion of the two connected masses and of the mass center proves completely non-Keplerian, beginning with the libration around local vertical due to the considerable residual kinetic moment at connection. A practical application of the quasi-vertical libration is in orbital passive debris collector, when a sandwich composite large panel is orbited for long periods of time for collecting small mass, high velocity Earth orbit debris. The most promising and controversial application of such long tethers resides in the anchoring technique to achieve the skeleton of a future space elevator. The stability of motion is an important aspect which is approached my numerical simulations.

An Evolutionary Algorithm Approach To Multi-Objective Scheduling Of Space Network Communications

We describe an evolutionary multi-objective optimization approach to the scheduling problem presented by the NASA's Deep Space Network (DSN). This network is the communications system that supports NASA and other space missions from high earth orbit to the outer planets. Today the DSN consists of a small number of large antennas, but in the future it is expected to incorporate several large arrays of smaller antennas that can be flexibly combined for each spacecraft communication session. Multi-objective techniques for schedule optimization have the attractive advantage of explicitly capturing the constraints and preferences of the missions that use the DSN, as well as those based on system-level considerations, and providing unique insight into trade-offs among competing requirements. We have investigated problem representation issues, objective and constraint formulations, and multi-objective optimization techniques that can be applied to this problem. We describe our initial results using an evolutionary algorithm on an illustrative sample problem (contention for a single antenna), and on a projected 2015 mission set with a three-site, 300 antenna array. The results are very promising, not only for generating initial schedules, but also for resolving conflicts where there is severe resource contention.

Relative Semimajor Axis Uncertainty in High Earth Orbits

S EVERAL proposed missions have identified satellite formation flying in high Earth orbit (HEO) as an enabling technology for increasing the science return while decreasing total mission risk [1,2]. Precision formationflying inHEO requires accurate estimation of the semimajor axis difference between the vehicles for state prediction, formation control, and formation maneuver planning [3– 5]. This Note derives the complete linearized mapping relating relative position and velocity error to semimajor axis uncertainty for HEO formations. The majority of relative navigation filters estimate the state of the secondary vehicle with respect to the primary in Cartesian position and velocity coordinates [6–13]. The uncertainty in the semimajor axis is found by rotating the estimated position and velocity covariance into orbital element space. Carpenter and Schiesser [3] presented the complete linearized relationship between absolute position and velocity uncertainty and semimajor axis error; however, the form of their result is not particularly intuitive. Carpenter and Alfriend [4] used a similar method to develop a simplified expression for the semimajor axis uncertainty that assumes the radial position and velocity and intrack velocity are the primary contributions. However, this assumption is not justified for eccentric orbits. The generalized relationship between relative position and velocity uncertainty and semimajor axis uncertainty is derived here explicitly. The resulting equation shows that the semimajor axis error is dependent on all of the coplanar position and velocity uncertainties in highly eccentric orbits. The coordinate frames that form the foundation of this research are described in Sec. II. The generalized relationship between position and velocity error and semimajor axis uncertainty is derived in Sec. III, and the conclusions are summarized in Sec. IV.

Earth-moon Trajectory Optimization Using Solar Electric Propulsion

The optimization of the Earth-moon trajectory using solar electric propulsion is presented. A feasible method is proposed to optimize the transfer trajectory starting from a low Earth circular orbit (500 km altitude) to a low lunar circular orbit (200 km altitude). Due to the use of low-thrust solar electric propulsion, the entire transfer trajectory consists of hundreds or even thousands of orbital revolutions around the Earth and the moon. The Earth-orbit ascending (from low Earth orbit to high Earth orbit) and lunar descending (from high lunar orbit to low lunar orbit) trajectories in the presence of J2 perturbations and shadowing effect are computed by an analytic orbital averaging technique. A direct/indirect method is used to optimize the control steering for the trans-lunar trajectory segment, a segment from a high Earth orbit to a high lunar orbit, with a fixed thrust-coast-thrust engine sequence. For the trans-lunar trajectory segment, the equations of motion are expressed in the inertial coordinates about the Earth and the moon using a set of nonsingular equinoctial elements inclusive of the gravitational forces of the sun, the Earth, and the moon. By way of the analytic orbital averaging technique and the direct/indirect method, the Earth-moon transfer problem is converted to a parameter optimization problem, and the entire transfer trajectory is formulated and optimized in the form of a single nonlinear optimization problem with a small number of variables and constraints. Finally, an example of an Earth-moon transfer trajectory using solar electric propulsion is demonstrated.

High-Altitude Satellite Relative Navigation Using Carrier-Phase Differential Global Positioning System Techniques

A new carrier-phase differential global positioning system relative navigation estimator has been developed that extends the use of carrier-phase differential global positioning system techniques to spacecraft formations that operate at geostationary altitudes and above. The estimator achieves rapid convergence to the carrier-phase ambiguities and incorporates a cycle slip detection and recovery algorithm. It solves a linearized problem using least-squares square-root information processing that does not require spacecraft dynamics models. In this context, integer ambiguities are resolved using an integer least-squares algorithm. The cycle slip algorithm identifies the slip channel by statistical hypothesis testing and estimates the magnitude of the slip. Global positioning system receiver-in-the-loop tests with simulated low-Earth orbit data show nearly instantaneous convergence to the correct integer ambiguities and relative position error magnitudes of less than 3 mm. Truth-model simulations are used to simulate geostationary orbits and high-Earth orbit scenarios. The geostationary orbit scenario produces nearly instantaneous convergence to the ambiguities and error magnitudes of less than 0.1 m. The high-Earth orbit case at a radial distance of 17.8 Earth radii converges in minutes with error magnitudes of less than 3 m. Cycle slips, present in the hardware-in-the-loop simulations, are detected and corrected without significant accuracy degradation.

Analysis of relative navigation in high earth orbits

Formation flying in high Earth orbit relies on accurate relative navigation information for both precise formation control and accurate interpretation of science data. A relative navigation filter is developed here that enhances the linearity of the dynamic model by framing the state in Keplerian elements. The dynamic model and filter performance are presented for medium and high Earth orbits with limited measurement availability. The effect of higher order gravity terms, solar radiation pressure, the third-body effects of the Sun and Moon, and atmospheric drag on estimation accuracy is considered. Relative semimajor axis accuracy of 0.81 m and relative position accuracy of 0.89 m are obtained for vehicles separated by 10–170 km using GPS measurements only. This represents an order of magnitude improvement over previously reported relative navigation performance in high Earth orbit.

Nonplanar LEO‐LEO aeroassisted space rendezvous

PurposeIn order to solve nonplanar LEO‐LEO aeroassisted space rendezvous, this paper aims to study an active phasing method based on orbital preliminary adjusting scheme out of atmosphere.Design/methodology/approachIn order to add atmospheric entry velocity, orbital preliminary adjusting out of atmosphere is presented and the orbital altitude of high earth orbit (HEO) is selected eclectically. Nonplanar HEO‐LEO aeroassisted orbital transfer problem is studied in detail. According to the standard atmospheric flight trajectory, the locations of deorbit points in HEO are determined and the standard phase angle between orbital transfer vehicle (OTV) in HEO and target in low earth orbit (LEO) is obtained so that OTV and target meet space rendezvous demand. Finally, the active phasing method is studied so that the standard phase angle can be satisfied when OTV is transferred to HEO. So space rendezvous can be realized under the help of aeroassisted orbital transfer technique once the standard phase angle is satisfied.FindingsNonplanar LEO‐LEO orbital transfer depending on entirely propulsive will use enormous fuel and the phasing problem will be most difficult in nonplanar LEO‐LEO space rendezvous mission. However, the fuel consumption can be saved and rendezvous mission can be finished in an advisable time when nonplanar LEO‐LEO aeroassisted orbital transfer technology is applied properly.Originality/valueAeroassisted space rendezvous method is presented in this paper. Orbital preliminary adjusting out of atmosphere is studied in order to add atmospheric entry velocity and the active phasing method for realizing space rendezvous is integrated in preliminary adjusting scheme.

Stabilization of Satellite Motion Relative to a Coulomb Spacecraft Formation

Coulomb spacecraft are satellites that can actively control their electrostatic charge and thus exploit intervehicle electrostatic forces to control tight relative motion. The stabilization problem of a cluster of unequal Coulomb spacecraft is studied. Previous research has developed a nonlinear control law to stabilize the relative motion of one satellite relative to another. With only two spacecraft present, with equal mass and charging limits, Newton’s second law greatly simplified the control development. This control strategy is generalized here to stabilize the relative motion of a satellite relative to a larger cluster of Coulomb spacecraft with unequal satellite masses and individual charge saturation limits. The chief cluster motion can be either circular or elliptic. The nonlinear control methodology exploits an orbit element difference description of the satellite relative motion. Although the control is shown to stabilize the relative motion of a Coulomb satellite about any set of desired orbit element differences, convergence is shown thus far only when controlling exclusively the semimajor axis differences. Thus, this control is able to achieve bounded relative motion of the Coulomb satellite, even in the presence of saturation constraints. A simple structured control approach is numerically investigated to control the entire cluster. Numerical simulations illustrate the relative motion control behavior with the cluster chief or center of mass being on an elliptic high Earth orbit.

Impact of Limitations in Geophysical Background Models on Follow-on Gravity Missions

Purpose of this article is to demonstrate the effect of background geophysical corrections on a follow-on gravity mission. We investigate the quality of two effects, tides and atmospheric pressure variations, which both act as a surface load on the lithosphere. In both cases direct gravitational attraction of the mass variations and the secondary potential caused by the deformation of the lithosphere are sensed by a gravity mission. In order to assess the current situation we have simulated GRACE range-rate errors which are caused by differences in present day tide and atmospheric pressure correction models. Both geophysical correction models are capable of generating range-rate errors up to 10 μm/s and affect the quality of the recovered temporal and static gravity fields. Unlike missions such as TOPEX/Poseidon where tides can be estimated with the altimeter, current gravity missions are only to some degree capable of resolving these (geo)physical limitations. One of the reasons is the use of high inclination low earth orbits without a repeating ground track strategy. The consequence is that we will face a contamination of the gravity solution, both in the static and the time variable part. In the conclusions of this paper we provide suggestions for improving this situation, in particular in view of follow-on gravity missions after GRACE and GOCE, which claim an improved capability of estimating temporal variations in the Earth’s gravity field.

Kronos: a satellite for astrotomography

Kronos, a Medium Explorer concept, is designed to initiate a new era in astrophysics, fully opening the domain of time by enabling extended simultaneous multiwavelength (X-ray, UV, and optical) spectroscopic monitoring of variable accretion-driven sources. The primary design goal of the mission is to obtain high-fidelity velocity-delay maps of the broad-line region in active galactic nuclei (AGNs) and thus determine the geometry and structure of AGNs on microarcsecond scales. Doppler tomography and eclipse mapping techniques will be used to characterize and map Galactic binary systems, revealing the details of the physics of accretion processes in black hole, neutron star, and white dwarf binary systems. The high-Earth orbit of Kronos enables well-sampled, high time-resolution observations, critical for the innovative and sophisticated methods that are used to understand the accretion flows, mass outflows, jets, and other phenomena found in accreting sources.

A new fast cloud propagator for use in the GEO regime

A new, fast debris cloud propagator is currently under development at the University of Southampton. This propagator, which is dedicated to the geostationary (GEO) regime, as well as other high-Earth orbits, propagates the debris cloud as a whole, rather than as a number of individual fragments or pseudo-fragments. This fast cloud propagator (FCP) first identifies the number distributions of debris fragments in six-dimensional element-space and also identifies the distributions of element relationships with one another; for example, the distribution of fragments in an eccentricity vs. inclination scatter plot. The modes and extremities of the distributions, as well as other carefully chosen distribution parameters, numbering 73 in total, are then propagated to the required epoch and the number distributions and element relationships are re-created from the propagated distribution parameters. The debris cloud is accurately re-created – and a considerable speed increase is gained over a conventional propagator, which typically evolves the large number of fragments found in the original debris cloud.

Study of Interspacecraft Coulomb Forces and Implications for Formation Flying

In the course of exploiting spacecraft formations for use in separated interferometry (or other missions), it is possible that the separation distance between vehicles will be on the order of 10 m. The effects of spacecraft charging on the dynamics of very closely spaced formations are investigated. For certain high-Earth orbits, the ambient plasma conditions will conspire to produce significant spacecraft charging in an environment with a plasma Debye length of more than 100 m. For such conditions, it is shown that the potential exists to develop disruptive interspacecraft Coulomb forces and torques, with magnitude comparable to candidate formation-keeping thrusters over distances of tens of meters. Because of the unexpectedly large interaction forces, the concept of purposely charging spacecraft to affect formation-keeping Coulomb forces is also explored. Analytic methods are developed that show the existence of static equilibrium formations in Earth orbit using only inter-vehicle coulomb forces for one-, two-, and three-dimensional formations. Such Coulomb formations would be free of the risk of plume contamination due to thrusters firing in close proximity. Figures of merit for the proposed Coulomb control system are calculated in a manner analogous to traditional propulsion systems, and it is shown that required forces can be created with milliwatts of power, can be controlled on a millisecond timescale, and imply specific impulses that can be as high as 1013 seconds.

Transfers to Earth centered orbits via lunar gravity assist

The use of lunar gravity assist to transfer payloads to Earth centered orbits is investigated. Such payloads include communications spacecraft bound for geostationary orbits or spacecraft bound for high inclination Earth orbits. Though payloads need to be injected on translunar trajectories with apogees at the lunar distance, the use of lunar gravity to change the inclination of the orbit provides (for a range of launch site latitudes) a substantial increase in the payload mass delivered or a reduction in the required characteristic velocity required to achieve orbit. By using the restricted three body problem as an approximate model for the motion of a spacecraft in the Earth–Moon system, a perturbation procedure and an optimization algorithm are used to construct nominal solutions.

Hubble Space Telescope at twelve years of age

The Hubble Space Telescope was deployed from the Space Shuttle Discovery into a 380-mile high Earth orbit on April 25, 1990. It subsequently made outstanding astronomical discoveries with its 8-foot (2.4-meter) telescope and other scientific instruments. Critical to the successful observations was continuous availability of power from its solar arrays during sunlit periods, and from nickel-hydrogen batteries when the satellite was in the Earth's shadow. The adopted nickel-hydrogen batteries were carefully selected and tested to confirm their depth-of-discharge and operating temperature that delivered the longest life in charge/discharge cycling service. These batteries had a design life of 7 years. At 12 years after launch the Hubble batteries have delivered more charge/discharge cycles than any other batteries in low-Earth orbit. However, the Hubble batteries have been subjected to many unexpected stresses, and peculiar reductions in battery capacity have been observed. Battery replacement requires a costly trip to the Hubble Space Telescope by astronauts, so the remaining useful life of the batteries must be predicted. Already in four servicing missions, astronauts have replaced or modified optics, solar arrays, a power control unit, and various science packages. A fifth servicing mission is scheduled in 2004. This paper discusses battery charging hardware and software controls, history of battery events in Hubble, cell performance model and spare battery tests, and capacity walkdown.

Performance characteristics of high resolution Compton telescopes

To build upon the goals of the upcoming INTEGRAL mission, the next generation soft γ -ray (0.2-20 MeV) observatory will require improved sensitivity to continuum and nuclear line emission, while maintaining high spectroscopy to resolve Doppler shifts and broadenings. We present the simulated performance of two germanium Compton telescope (GCT) designs in high Earth orbit conditions. The goal of this work is both to determine realistic sensitivities for the GCTs, as well as determine the geometry which best takes advantage of the spectral, imaging, and stopping power of germanium detectors. We find that a novel Compact design optimizes the sensitivity to both nuclear line and continuum sources, with an effective factor of 2.0 improvement over a Classical design. The Compact GCT would provide an order of magnitude improvement in sensitivity over INTEGRAL.

GPS Receiver Architecture and Expected Performance for Autonomous Navigation in High Earth Orbits

ABSTRACT: GPS is an enabling technology for autonomous spacecraft navigation, including mission concepts involving formation flying in highly eccentric orbits. However, to date most spaceborne GPS applications have been limited to low earth orbits or postprocessing applications. This paper describes the problems associated with using GPS in the high earth orbit (HEO) environment and presents a GPS receiver design optimized for these applications. The design builds upon the PiVoT receiver developed by Goddard Space Flight Center by incorporating the GPS Enhanced Orbit Determination (GEODE) navigation filter, a high-quality clock, and enhancements to basic receiver acquisition and tracking algorithms. Predicted navigation performance for several representative missions, including geosynchronous and geosynchronous transfer orbits, is presented. The analysis shows that a modest reduction in the acquisition threshold of the receiver enables reliable navigation performance even in very high orbits with unfavorable antenna placement restrictions.

Techniques for galactic dust measurements in the heliosphere

Galactic interstellar dust (ISD) is the major ingredient in planetary formation. However, information on this important material has been extremely limited. Recently, the Ulysses dust detector has identified and measured interstellar dust outside 1.8 AU from the Sun at ecliptic latitudes above 50°. Inside this distance it could not reliably distinguish interstellar from interplanetary dust. Modeling the Ulysses data suggests that up to 30% of dust flux with masses above 10−16 kg at 1 AU is of interstellar origin. From the Hiten satellite in high eccentric orbit about the Earth, there are indications that ISD indeed reaches the Earth's orbit. Two new missions carrying dust detectors, Cassini and Stardust, will greatly increase our observational knowledge. In this paper we briefly review instruments used on these missions and compare their capabilities. The Stardust mission [Brownlee et al., 1996] will analyze the local interstellar dust population by an in situ chemical analyzer and collect ISD between 2 and 3 AU from the Sun. The dust analyzer on the Cassini mission will determine the interstellar dust flux outside Venus' orbit and will provide also some compositional information. Techniques to identify the ISD flux levels at 1 AU are described that can quantify the interstellar dust flux in high Earth orbit (outside the debris belts) and provide chemical composition information of galactic dust.

Internal charging: a preliminary environmental specification for satellites

Internal charging has been indicated as the cause of many satellite anomalies. In some cases, it has been argued that internal charging has caused satellite system failures. It may be possible to predict the occurrence of an internal charging threat in the future but that does not remove the necessity to know what the environmental threat may be. The existence, for many years now, of energetic electron and internal charging measurements in the inner magnetosphere provides the ability to identify the threat levels and generate internal charging specifications. The specifications must embody the worst-case environments that can be expected from magnetic storms and extreme solar/interplanetary conditions. Internal-charging environment specifications are needed by satellite manufacturers for setting design requirements for their systems. The specifications are also required for defining the worst-case energetic electron flux levels and total fluence levels from events. These are used for electron beam testing to verify that critical systems are immune to internal charging, for performing tests on subsystems, and for use in anomaly investigation programs. Data from the GOES, LANL, and CRRES satellites were used to develop preliminary internal-charging environment specifications for some commonly used orbits such as geosynchronous and Molniya or high earth orbits (HEO), and they are discussed. A specification is also given for a highly elliptical equatorial orbit that is used for lunar-transfer orbital maneuvers. The preliminary environments were used to estimate the shielding required to reduce worst-case electron fluxes to safe levels for these orbits.

Asteroidal defence

Continuing this week's asteroidal theme, Daedalus has a plan for what we could do should a stray asteroid or comet smash into the Earth. His idea is that a defence system could be placed in a high Earth orbit. But this system would not consist of nuclear missiles -- rather, it would be another asteroid.