Design of optimal earth pole-sitter transfers using low-thrust propulsion

Recent studies have shown the feasibility of an Earth pole-sitter mission using low-thrust propulsion. This mission concept involves a spacecraft following the Earth's polar axis to have a continuous, hemispherical view of one of the Earth's poles. Such a view will enhance future Earth observation and telecommunications for high latitude and polar regions. To assess the accessibility of the pole-sitter orbit, this paper investigates optimum Earth pole-sitter transfers employing low-thrust propulsion. A launch from low Earth orbit (LEO) by a Soyuz Fregat upper stage is assumed after which solar electric propulsion is used to transfer the spacecraft to the pole-sitter orbit. The objective is to minimize the mass in LEO for a given spacecraft mass to be inserted into the pole-sitter orbit. The results are compared with a ballistic transfer that exploits manifold-like trajectories that wind onto the pole-sitter orbit. It is shown that, with respect to the ballistic case, low-thrust propulsion can achieve significant mass savings in excess of 200 kg for a pole-sitter spacecraft of 1000 kg upon insertion. To finally obtain a full low-thrust transfer from LEO up to the pole-sitter orbit, the Fregat launch is replaced by a low-thrust, minimum time spiral, which provides further mass savings, but at the cost of an increased time of flight.     

Rosetta enters hibernation

The International Rosetta Mission was launched on 2nd March 2004 on its 10 years journey to comet 67P/Churyumov–Gerasimenko. Rosetta will reach the comet in 2014, orbit it for about 1.5 years down to distances of a few kilometres and deliver the Lander Philae onto its surface.Following the fly-by of Asteroid (21-)Lutetia in 2010, Rosetta continued its travel towards the planned comet encounter in 2014. In this phase Rosetta became the solar-powered spacecraft that reached the largest Sun distances in history of spaceflight, up to an aphelion at 5.3 AU in October 2012. At distances above 4.5 AU the spacecraft's solar generator power is not sufficient to keep all spacecraft systems active. Therefore in June 2011 the spacecraft was spun up to provide gyroscopic stabilisation, and most of its on-board units, including those used for attitude control and communications, were switched off. Over this “hibernation” phase of about 2.5 years the spacecraft will keep a minimum of autonomy active to ensure maintenance of safe thermal conditions.After Lutetia fly-by, flight controllers had to tackle two anomalies that had significant impacts on the mission operations. A leak in the reaction control subsystem was confirmed and led to the re-definition of the operational strategy to perform the comet rendezvous manoeuvres planned for 2011 and 2014. Anomalous jumps detected in the estimated friction torque of two of the four reaction wheels used for attitude control forced the rapid adoption of measures to slow down the wheels degradation. This included in-flight re-lubrication activities and changes in the wheels operational speed regime.Once the troubleshooting of the two anomalies was completed, and the related operational scenarios were implemented, the first large (790 m/s) comet rendezvous manoeuvre was executed, split into several long burns in January and February 2011. The second burn was unexpectedly interrupted due to the anomalous behaviour of two thrusters, causing attitude off-pointing. Flight controllers modified the thrusters operation parameters in the on-board software and managed to re-start the sequence of burns and successfully complete the manoeuvre. After the manoeuvre, preparation for the critical spin-up and hibernation entry activities, planned for June 2011, began.This paper presents the activities carried out on Rosetta in the final year before hibernation entry. The major anomalies and the related troubleshooting and workaround solutions are detailed. Lessons learned from the operation of the first spacecraft operating with solar power at Jupiter-like distances from the Sun are presented and discussed.     

Analytical solutions for the relative motion of spacecraft subject to Lorentz-force perturbations

A spacecraft capable of producing higher-than-natural electrostatic charges may achieve propellantless orbital maneuvering via the Lorentz-force interaction with a planetary magnetic field. Development of maneuver strategies for these propellantless vehicles is complicated by the fact that the perturbative Lorentz force acts along only a single line of action at any instant. Relative-motion dynamical models are developed that lead to approximate analytical solutions for the motion of charged spacecraft subject to the Lorentz force. These solutions indicate that the principal effects of the Lorentz force on a spacecraft in a circular orbit are to change the intrack position and to change the orbit plane. A rendezvous example is presented in which a spacecraft with a specific charge of ?3.81 $\times$ 10^?4 C/kg reaches a target vehicle initially 10 km away (on the same equatorial low-Earth orbit) in 1 day. Fly-around maneuvers may be achieved in low-Earth orbit with specific charges on the order of 0.001 C/kg.     

Compensation of gravitational attraction disturbance to pure gravity orbit for Inner Formation Flying System

The Inner Formation Flying System (IFFS) consisting of an outer satellite and an inner satellite which is a solid sphere proof mass freely flying in the shield cavity can construct a pure gravity orbit to precisely measure the earth gravity field. The gravitational attraction on the inner satellite due to the outer satellite is a significant disturbance source to the pure gravity orbit and is required to be limited to 10⁻¹¹ m s⁻² order. However, the gravitational disturbance force was on 10⁻⁹ m s⁻² order actually and must be reduced by dedicated compensation mass blocks. The region of relative motion of the inner satellite about its nominal position is within 1 cm in dimension, which raises the complexity of the compensation blocks design. The iterative design strategy of the compensation blocks based on reducing the gravitational attraction at the nominal position of the inner satellite is presented, aiming to guarantee the gravitational force in the relative motion region within requirements after the compensation. The compensation blocks are designed according to the current status of IFFS, and the gravitational disturbance force in the region is reduced to 10⁻¹¹ ms⁻² order with minimized adding mass.     

Design and on-orbit performance of the attitude determination and control system for the ZDPS-1A pico-satellite

The ZDPS-1A pico-satellite, developed by the Zhejiang University, is featured with a three-axis stabilizing capability. It is 15×15×15 cm³ cube-shaped satellite with a total mass of 3.5 kg. ZDPS-1A is the first pico-satellite that has been launched successfully in China. The mission of ZDPS-1A is on-orbit system verification of student-build pico-satellite and wide range earth observation with a micro panoramic camera. A miniature momentum wheel is employed to offer gyro stiffness stability in the pitch (orbit normal) axis. Magnetic coils are employed to generate control torques to achieve the three-axis stabilization of nadir-pointing. The attitude sensors employed in the design include two three-axis magnetometers (TAMs), a three-axis gyro, and two sun sensors. Both ground simulations and on-orbit testing are conducted to verify the feasibility of the given attitude determination and control system (ADCS).     

Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

A new and innovative type of gridded ion thruster, the “Dual-Stage 4-Grid” or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed.A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a “medium-term” SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific impulse is lower at 3500 s which is well within conventional gridded ion thruster capability.     

Rosetta visits asteroid (21)Lutetia

The International Rosetta Mission, cornerstone of the European Space Agency Scientific Programme, was launched on 2nd March 2004 to its 10 years journey to comet Churyumov–Gerasimenko. Rosetta will reach the comet in summer 2014, orbit it for about 1.5 years down to distances of a few Kilometres and deliver the Lander Philae onto its surface. After its successful asteroid fly-by in September 2008, Rosetta came back to Earth, for the final gravity acceleration towards its longest heliocentric orbit, up to a distance of 5.3 AU. It is during this phase that Rosetta crossed for the second time the main asteroids belt and performed a close encounter with asteroid (21)Lutetia on the 10th of July 2010 at a distance of ca. 3160 km and a relative velocity of 15 km/s. The payload complement of the spacecraft was activated to perform highly valuable scientific observations. The approach phase to the celestial body required a careful and accurate optical navigation campaign that will prove to be useful also for the comet approach phase. The experience gained with first asteroid flyby in 2008 was fed back into the operations definition and preparation for this highly critical phase; this concerns in particular the operations of the navigation camera for the close-loop autonomous asteroid tracking and of the main scientific camera for high resolution imaging. It was shortly after the flyby that Rosetta became the solar-powered spacecraft to have flown furthest from the Sun (>2.72 AU). This paper presents the activities carried out and planned for the definition, preparation and implementation of the asteroid flyby mission operations, including the test campaign conducted to improve the performance of the spacecraft and payload compared to the previous flyby. The results of the flyby itself are presented, with the operations implemented, the achieved performance and the lessons learned.     

Super-capacitor energy storage for micro-satellites: Feasibility and potential mission applications

Small satellites, weighting between 100 and 200 kg, have witnessed increasing use for a variety of space applications including remote sensing constellations and technology demonstrations. The energy storage/stored power demands of most spacecraft, including small satellites, are currently accommodated by rechargeable batteries—typically nickel–cadmium cells (specific energy of 50 Wh kg⁻¹), or more recently lithium-ion cells (150 Wh kg⁻¹). High energy density is a primary concern for spacecraft energy storage design, and these batteries have been sufficient for most applications. However, constraints on the allowable on-board battery size have limited peak power performance such that the maximum power supply capability of small satellites currently ranges between only 70 and 200 W. This relatively low maximum power limits the capabilities of small satellites in terms of payload design and selection. In order to enhance these satellites' power performance, the research reported in this paper focused on the implementation of super-capacitors as practical rechargeable energy storage medium, and as an alternative to chemical batteries. Compared to batteries, some super-capacitors are able to supply high power at high energy-efficiency, but unfortunately they still have a very low energy density (5–30 Wh kg⁻¹). However, the provision of this high power capability would considerably widen the range of small satellite applications.     

Measured surface temperatures of the Hayabusa capsule during re-entry determined from ground observation

The Hayabusa sample return capsule, which contained asteroid samples, re-entered the Earth's atmosphere on June 13, 2010. An ablative carbon-phenolic thermal protection system (TPS) was used to enable a safe return for the small capsule and the containing samples. Besides a research aircraft operated by NASA with a wide range of imaging and spectrographic cameras for remote sensing of the radiation of the Hayabusa capsule during its entry flight, observation from ground based stations has been realized. We participated in the ground based observation campaign with two instruments for spectroscopic and photometric measurements aiming to detect the surface temperature and the plasma radiation in front of the re-entering capsule. The system consists in an infrared camera and a wide range miniature fibre spectrometer. The paper presents the setup, the laboratory calibration procedure, and correction for transmission. The surface temperature of the capsule reached a peak of 3250 K when the capsule was at an altitude of 55.95 km. The thermographic camera measures independently slightly higher temperature at peak heating (3308 K).     

Field emission performance of multiwalled carbon nanotubes for a low-power spacecraft neutraliser

Field electron emission from aligned multiwalled carbon nanotubes has been assessed to determine if the performance, defined by power consumption, lifetime and emission current, is suitable for use in spacecraft charge neutralisation for field emission electric propulsion (FEEP). Carbon nanotubes grown by chemical vapour deposition (CVD) were mounted on a dual in line chip with a macroscopic (nickel mesh) extractor electrode mounted ～1 mm above the tubes. The nanotubes’ field emission characteristics (emission currents, electron losses and operating voltage) were measured at ～10^?4 Pa. An endurance test of one sample, running at a software-controlled constant emission current lasted >1400 h, approaching the longest known FEEP thruster lifetime. The emission corresponds to a current density of ～10 mA/cm² at a voltage of 150 V. These results, implementing mature extractor-electrode geometry, indicate that carbon nanotubes have considerable potential for development as robust, low-power, long-lived electron emitters for use in space.     

Cassini Saturn-escape trajectories to Jupiter,Uranus, and Neptune

Potential encore-mission scenarios have been considered for the Cassini mission. In this paper we discuss one of the end-of-life scenarios in which the Cassini spacecraft could perform a Saturn escape via gravity assists from Titan. It is shown that such satellite-aided escape requires a small deterministic maneuver (e.g., Δv<50 m/s), but provides enough energy for the Cassini spacecraft to reach a range of targets in our Solar System, as close to the Sun as the asteroid belt or as far as the Kuiper belt. The escape sequence could be initiated from an arbitrary point during the on-going Cassini mission. Example tours are presented in which the final Titan flyby places the spacecraft into ballistic trajectories that reach Jupiter, Uranus, and Neptune. After years of heliocentric flight, the spacecraft could impact on the target gas giant or perform a flyby to escape from the Solar System (if not to another destination). The concept can be generalized to a new kind of missions, including nested-grand tours, which may involve satellite-aided captures and escapes at more than one planet.     

Parabolic maneuvers of the Swiss Air Force fighter jet F-5E as a research platform for cell culture experiments in microgravity

Long-term sensitivity of human cells to reduced gravity has been supposed since the first Apollo missions and was demonstrated during several space missions in the past. However, little information is available on primary and rapid gravi-responsive elements in mammalian cells. In search of rapid-responsive molecular alterations in mammalian cells, short-term microgravity provided by parabolic flight maneuvers is an ideal way to elucidate such initial and primary effects. Modern biomedical research at the cellular and molecular level requires frequent repetition of experiments that are usually performed in sequences of experiments and analyses. Therefore, a research platform on Earth providing frequent, easy and repeated access to real microgravity for cell culture experiments is strongly desired. For this reason, we developed a research platform onboard the military fighter jet aircraft Northrop F-5E “Tiger II”. The experimental system consists of a programmable and automatically operated system composed of six individual experiment modules, placed in the front compartment, which work completely independent of the aircraft systems. Signal transduction pathways in cultured human cells can be investigated after the addition of an activator solution at the onset of microgravity and a fixative or lysis buffer after termination of microgravity. Before the beginning of a regular military training flight, a parabolic maneuver was executed. After a 1 g control phase, the parabolic maneuver starts at 13,000 ft and at Mach 0.99 airspeed, where a 22 s climb with an acceleration of 2.5g is initiated, following a free-fall ballistic Keplerian trajectory lasting 45 s with an apogee of 27,000 ft at Mach 0.4 airspeed. Temperature, pressure and acceleration are monitored constantly during the entire flight. Cells and activator solutions are kept at 37 °C during the entire experiment until the fixative has been added. The parabolic flight profile provides up to 45 s of microgravity at a quality of 0.05g in all axes. Access time is 30 min before take-off; retrieval time is 30 min after landing. We conclude that using military fighter jets for microgravity research is a valuable tool for frequent and repeated cell culture experiments and therefore for state-of-the art method of biomedical research.     

Flight performance analysis of GRACE K-band ranging instrument with simulation data

A dual one-way ranging (DOWR) system provides very high accuracy range measurements between two satellites. The GRACE satellite mission implements the DOWR, called KBR (K-band ranging), to measure very small inter-satellite range change in order to map the Earth gravity field. The flight performance of the KBR is analyzed by using a hybrid software simulator that incorporates actual satellite orbit data into a comprehensive KBR simulator, which was earlier used for computing the GRACE baseline accuracy. Three types of experiments were performed. First is the comparison of the flight data with the simulated data in spectral domain. Second is the comparison of double differenced noise level. Third is the comparison of the range-rate difference with GPS clock estimates. The analysis shows a good agreement with the simulation model except some excessive high frequency noise, e.g. 10⁻⁴ m/√Hz at 0.1 Hz. The range-rate difference shows 0.003 cyc/s discrepancy with the clock estimates. These analyses are helpful to refine the DOWR simulation model and can be benefit to future DOWR instrument development.     

Launch and early operations of Herschel and Planck

On 14 May 2009 the European Space Agency launched 2 space observatories: Herschel (with a 3.5 m mirror it is the largest space telescope ever) will collect long-wavelength infrared radiation and will be the only space observatory to cover the spectral range from far-infrared to sub-millimetre wavelengths, and Planck will look back at the dawn of time, close to the Big Bang, and will examine the Cosmic Microwave Background (CMB) radiation to a sensitivity, angular resolution and frequency range never achieved before. This paper will present the Flight Dynamics, mission analysis challenges and flight results from the first 3 months of these missions.Both satellites were launched on the same Ariane 5 and travelled to the L2 Lagrange point of the sun–earth system 1.5 million km from the earth in the opposite direction of the sun. There they were injected to a quasi-halo orbit (Herschel) with the dimension of typically 750,000 km×450,000 km, and a Lissajous orbit (Planck) of 300,000 km×300,000 km.In order to reach these Lissajous orbits it is mandatory to perform large trajectory correction manoeuvres during the first days of the mission. Herschel had its main manoeuvres on the first day. Planck had to be navigated on the first day and by a mid-course correction manoeuvre, the L2 orbit insertion manoeuvre was planned on day 50. If these slots were missed, fuel penalties would rapidly increase.This posed a heavy load on the operations teams because both spacecrafts have to be thoroughly checked out and put into the correct modes of their attitude control systems during the first hours after launch.The sequence of events will be presented and explained and the orbit determination results as well as the manoeuvre planning will be emphasised.     

Dose estimates in a lunar shelter with regolith shielding

Beyond the Earth's atmosphere, galactic cosmic radiation (GCR) and solar energetic particles (SEPs) are a significant hazard to both manned and robotic missions. For long human missions on the lunar surface (months to a year) a radiation shelter is needed for dose mitigation and emergency protection in case of solar events. This paper investigates the interaction of source protons of solar events like those of February 1956 that emitted many fewer particles with energies up to 1000 MeV and of the October 1989 event of lower protons energy but higher fluence, with the lunar regolith and aluminum shielding of a lunar shelter. The shelter is 5 m in diameter and has a footprint of 5×8 m and a 10 cm thick aluminum support structure, however, actual thickness could be much smaller (～1–2 cm) depending on the weight of the regolith shielding piled on top. The regolith is shown to be slightly more effective than aluminum. Thus, the current results are still applicable for a thinner aluminum structure and increased equivalent (or same mass) thickness of the regolith. The shielding thicknesses to reduce the dose solely due to solar protons in the lunar shelter below those recommended by NASA to astronauts for 30 day-operation in space (250 mSv) and for radiation workers (50 mSv) are determined and compared. The relative attenuation of incident solar protons with regolith shielding and the dose estimates inside the shelter are calculated for center seeking, planar, and isotropic incidence of the source protons. With the center seeking incidence, the dose estimates are the highest, followed by those with isotropic incidence, and the lowest are those with the planar incidence.     

MESSENGER at Mercury: Early orbital operations

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft, launched in August 2004 under NASA's Discovery Program, was inserted into orbit about the planet Mercury in March 2011. MESSENGER's three flybys of Mercury in 2008–2009 marked the first spacecraft visits to the innermost planet since the Mariner 10 flybys in 1974–1975. The unprecedented orbital operations are yielding new insights into the nature and evolution of Mercury. The scientific questions that frame the MESSENGER mission led to the mission measurement objectives to be achieved by the seven payload instruments and the radio science experiment. Interweaving the full set of required orbital observations in a manner that maximizes the opportunity to satisfy all mission objectives and yet meet stringent spacecraft pointing and thermal constraints was a complex optimization problem that was solved with a software tool that simulates science observations and tracks progress toward meeting each objective. The final orbital observation plan, the outcome of that optimization process, meets all mission objectives. MESSENGER's Mercury Dual Imaging System is acquiring a global monochromatic image mosaic at better than 90% coverage and at least 250 m average resolution, a global color image mosaic at better than 90% coverage and at least 1 km average resolution, and global stereo imaging at better than 80% coverage and at least 250 m average resolution. Higher-resolution images are also being acquired of targeted areas. The elemental remote sensing instruments, including the Gamma-Ray and Neutron Spectrometer and the X-Ray Spectrometer, are being operated nearly continuously and will establish the average surface abundances of most major elements. The Visible and Infrared Spectrograph channel of MESSENGER's Mercury Atmospheric and Surface Composition Spectrometer is acquiring a global map of spectral reflectance from 300 to 1450 nm wavelength at a range of incidence and emission angles. Targeted areas have been selected for spectral coverage into the ultraviolet with the Ultraviolet and Visible Spectrometer (UVVS). MESSENGER's Mercury Laser Altimeter is acquiring topographic profiles when the slant range to Mercury's surface is less than 1800 km, encompassing latitudes from 20°S to the north pole. Topography over the remainder of the southern hemisphere will be derived from stereo imaging, radio occultations, and limb profiles. MESSENGER's radio science experiment is determining Mercury's gravity field from Doppler signals acquired during frequent downlinks. MESSENGER's Magnetometer is measuring the vector magnetic field both within Mercury's magnetosphere and in Mercury's solar wind environment at an instrument sampling rate of up to 20 samples/s. The UVVS is determining the three-dimensional, time-dependent distribution of Mercury's exospheric neutral and ionic species via their emission lines. During each spacecraft orbit, the Energetic Particle Spectrometer measures energetic electrons and ions, and the Fast Imaging Plasma Spectrometer measures the energies and mass per charge of thermal plasma components, both within Mercury's magnetosphere and in Mercury's solar-wind environment. The primary mission observation sequence will continue for one Earth year, until March 2012. An extended mission, currently under discussion with NASA, would add a second year of orbital observations targeting a set of focused follow-on questions that build on observations to date and take advantage of the more active Sun expected during 2012–2013. MESSENGER's total primary mission cost, projected at $446 M in real-year dollars, is comparable to that of Mariner 10 after adjustment for inflation.     

基于磁悬浮控制力矩陀螺的航天器姿态高精度高带宽测量方法

针对航天器姿态测量精度和带宽之间相互制约问题,提出一种基于磁悬浮陀螺的航天器姿态高精度、高带宽测量方法。根据刚体动力学和坐标变换原理建立磁悬浮转子径向转动合外力矩模型。在框架静止条件下,通过实时检测磁悬浮控制力矩陀螺（MSCMG）中的磁轴承电流、磁悬浮转子位移,计算出磁悬浮转子径向转动所受合外力矩以及磁悬浮转子径向偏转信息,间接得到航天器运动对磁悬浮转子径向转动作用力矩,进而求出航天器单轴姿态角速度和姿态角加速度。不同带宽下的仿真结果表明,本测量方法能同时检测出航天器单方向的姿态角速度和角加速度,并且满足高精度高带宽要求。     

A harness for enhanced comfort and loading during treadmill exercise in space

IntroductionLocomotor and some resistance exercises in space require a gravity replacement force in order to allow 1g-like ground reaction forces to be generated. Currently bungee cords, or other loading devices, interface with the crew member through a harness with a waist belt and shoulder straps. Crew members often find the application of the required loads to be uncomfortable, particularly at the hips.MethodsAn experimental harness was built that differed from previous in-flight designs by having a wider, moldable waist belt and contoured shoulder straps with additional padding. Eight subjects ran at 100% body weight (BW) loading for a total duration of 30 min per day on 12 days over a 3-week period in simulated 0-g conditions using horizontal suspension. A 100 mm Visual Analog Scale (VAS)¹ was used to assess harness-related and lower extremity discomfort at the end of each run.ResultsThe overall rating of harness discomfort decreased from 27 mm on the 100 mm scale on day 1 to 10 mm on day 12, with significant decreases recorded for the back and hip regions as well as the overall harness.DiscussionThe experimental harness allows for repeated exposure to 30-minute bouts of 100% BW loaded simulated 0-g running with levels of discomfort less than 30 mm on a VAS scale of 0–100 mm. We believe that the use of such a harness during on-orbit exercise countermeasures may allow exercise to be performed at levels which are more effective in preventing bone and muscle loss.     

Attitude pointing enhancement for combined energy and attitude control system

This paper extends the previous works that appeared in Acta Astronautica. An approach that incorporates the Active Force Control (AFC) technique into a conventional Proportional-Derivative (PD) controller is proposed for a 50 kg small satellite. Numerical treatments are performed to validate the effectiveness of AFC. The attitude control capability of the combined energy and attitude control system (CEACS) is expected to improve. The result shows an important attitude pointing enhancement for the CEACS attitude control task.     

Effects of uncertainties and flexible dynamic contributions on the control of a spacecraft full-coupled model

One of the most important problems for performing a good design of the spacecraft attitude control law is connected to its robustness when some uncertainty parameters are present on the inertial and/or on the elastic characteristics of a satellite. These uncertainties are generally intrinsic on the modeling of complex structures and in the case of large flexible structures they can be also attributed to secondary effects associated to the elasticity. One of the most interesting issues in modeling large flexible space structures is associated to the evaluation of the inertia tensor which in general depends not only on the geometric ‘fixed’ characteristic of the satellite but also on its elastic displacements which of course in turn modify the ‘shape’ of the satellite. Usually these terms can be considered of a second order of magnitude if compared with the ones associated to the rigid part of a structure. However the increasing demand on the dimension of satellites due to the presence for instance of very large solar arrays (necessary to generate power) and/or large antennas has the necessity to investigate their effects on their global dynamic behavior in more details as a consequence. In the present paper a methodology based on classical Lagrangian approach coupled with a standard Finite Element tool has been used to derive the full dynamic equations of an orbiting flexible satellite under the actions of gravity, gravity gradient forces and attitude control. A particular attention has been paid to the study of the effects of flexibility on the inertial terms of the spacecraft which, as well known, influence its attitude dynamic behavior. Furthermore the effects of the attitude control authority and its robustness to the uncertainties on inertial and elastic parameters has been investigated and discussed.