FLIGHT DYNAMICS CHALLENGES OF THE GERMAN ON-ORBIT SERVICING MISSION DEOS T. Rupp, T. Boge, R. Kiehling, F. Sellmaier Deutsches Zentrum für Luft- und Raumfahrt (DLR), German Space Operations Center (GSOC) Münchner Str. 20, 82234 Wessling, Germany, +49(8153) 28-2476 [email protected], [email protected], [email protected], [email protected]ABSTRACT On-orbit servicing missions build a novel class of spacecraft missions. In such kind of missions a spacecraft (servicer) is launched into an orbit of another spacecraft (client). The servicer is navigated to the client with the intention to manipulate it in a predefined manner. Both approach and coupling of two spacecraft in orbit lead to various challenges in the area of space flight dynamics. This paper describes such a scenario based upon a phase-A study of the envisaged German on-orbit servicing mission DEOS. 1. INTRODUCTION Numerous scientific and commercial satellite missions as well as various space shuttle and MIR station payload missions besides the present ISS-Columbus laboratory have been operated by DLR’s German Space Operations Center (GSOC) during the past 40 years. Ground control activities were mostly directed to a single spacecraft or an extensive but well determined payload segment. In the upcoming years, significant changes especially in the profile of satellite missions have to be expected. Future mission architectures will comprise more than one spacecraft with the demand to control the individual units in a coordinated manner. One approach to enable novel applications in space is to consider robotic on-orbit servicing and inspection missions, where a servicing spacecraft manipulates a client satellite in a pre-defined manner. This provides the opportunity to get away from the disposable character of today’s spacecraft towards reuse or re- configuration and will thus help to better preserve our assets in space. It ought to be mentioned here that satellite formations and constellations are a related class of future spacecraft missions. They have large similarities with on-orbit servicing missions and are recognized also as a means to achieve new scientific and technological objectives that could not be realized with a single satellite or uncoordinated individual ones. Increasing complexity and costs of satellite missions promote the idea looking for opportunities to extend the operational lifetime or improve the performance of a spacecraft instead of simply replacing it by a new one. Satellites in orbit can severely be affected by aging or degradation of its components and systems as well as by consumption of available resources. Also the disposal of spacecraft after the end of lifetime will play more and more an important role in the future, especially, if satellite orbits are of strategic importance. Presently DLR is supporting two spacecraft missions objected to on-orbit servicing tasks: The DEOS mission - DEutsche Orbitale Servicing Mission, a technology demonstration mission in low earth orbit, where various scenarios in the area of rendezvous and docking as well as re-entry capabilities will be considered. Second, the OLEV mission - Orbital Life Extension Vehicle, a mission designed to extend the lifetime of geostationary communications satellites.
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FLIGHT DYNAMICS CHALLENGES OF THE GERMAN ON-ORBIT SERVICING MISSION DEOS
T. Rupp, T. Boge, R. Kiehling, F. Sellmaier
Deutsches Zentrum für Luft- und Raumfahrt (DLR), German Space Operations Center (GSOC)
Münchner Str. 20, 82234 Wessling, Germany, +49(8153) 28-2476
[email protected], [email protected], [email protected], [email protected] ABSTRACT On-orbit servicing missions build a novel class of spacecraft missions. In such kind of missions a spacecraft (servicer) is launched into an orbit of another spacecraft (client). The servicer is navigated to the client with the intention to manipulate it in a predefined manner. Both approach and coupling of two spacecraft in orbit lead to various challenges in the area of space flight dynamics. This paper describes such a scenario based upon a phase-A study of the envisaged German on-orbit servicing mission DEOS.
1. INTRODUCTION Numerous scientific and commercial satellite missions as well as various space shuttle and MIR station payload missions besides the present ISS-Columbus laboratory have been operated by DLR’s German Space Operations Center (GSOC) during the past 40 years. Ground control activities were mostly directed to a single spacecraft or an extensive but well determined payload segment. In the upcoming years, significant changes especially in the profile of satellite missions have to be expected. Future mission architectures will comprise more than one spacecraft with the demand to control the individual units in a coordinated manner. One approach to enable novel applications in space is to consider robotic on-orbit servicing and inspection missions, where a servicing spacecraft manipulates a client satellite in a pre-defined manner. This provides the opportunity to get away from the disposable character of today’s spacecraft towards reuse or re-configuration and will thus help to better preserve our assets in space. It ought to be mentioned here that satellite formations and constellations are a related class of future spacecraft missions. They have large similarities with on-orbit servicing missions and are recognized also as a means to achieve new scientific and technological objectives that could not be realized with a single satellite or uncoordinated individual ones. Increasing complexity and costs of satellite missions promote the idea looking for opportunities to extend the operational lifetime or improve the performance of a spacecraft instead of simply replacing it by a new one. Satellites in orbit can severely be affected by aging or degradation of its components and systems as well as by consumption of available resources. Also the disposal of spacecraft after the end of lifetime will play more and more an important role in the future, especially, if satellite orbits are of strategic importance. Presently DLR is supporting two spacecraft missions objected to on-orbit servicing tasks: The DEOS mission - DEutsche Orbitale Servicing Mission, a technology demonstration mission in low earth orbit, where various scenarios in the area of rendezvous and docking as well as re-entry capabilities will be considered. Second, the OLEV mission - Orbital Life Extension Vehicle, a mission designed to extend the lifetime of geostationary communications satellites.
According to the developed phase-A study approach DEOS will consist of both a servicing and a dedicated client spacecraft, which will be launched together into an initial orbit. Primary mission goal is the capturing of a tumbling and non-cooperative client satellite with a manipulator on the servicing spacecraft and the re-entry (de-orbit) of the rigidly coupled configuration within a pre-defined orbit corridor. To achieve the envisaged goal a dedicated set of experiments has to be conducted where in general the complexity of the experiment execution will be stepwise increased over the mission period:
1. Far range formation flying experiments between servicing and client spacecraft have to be performed. This mission phase is characterized by methods of absolute navigation based on conventional GPS sensors and angle measurements provided by different ground stations. In addition the identification of dynamical parameters of the individual spacecrafts will be done during this phase.
2. During the rendezvous phases the following experiments will be performed:
- Approach of the servicing spacecraft to the non-cooperative client - Departure from the client - Execution of fly around and inspection maneuvers All these experiments are characterized by methods of relative navigation between servicing and client spacecraft, which are based on optical cameras or LIDAR sensors. They will be repeated several times under different illumination conditions.
3. The performance of the docking and berthing procedures between the servicing and the client spacecraft has to be demonstrated by different experiments. The berthing phase is characterized by grappling the client satellite using the so called manipulator end-effector of the servicing spacecraft and latching the client onto the berthing port. During the docking phase the servicing spacecraft approaches the client, inserts the docking interface of the client into the servicer’s docking port and latches the client.
4. Finally, different flight maneuvers will be performed in a rigidly coupled configuration.
They comprise the execution of combined attitude and orbit maneuvers, the identification of dynamical parameters of the coupled configuration and the execution of special on-orbit servicing tasks with respect to the client. At last de-orbiting of the configuration rigidly coupled by the manipulator arm is foreseen, executed as a purposive re-entry within a given re-entry corridor.
In addition to these specific flight dynamics tasks during the operational mission phases an extensive validation and verification program has to be performed during the various development steps on ground. Because the rendezvous and docking processes are the most critical mission phases the corresponding maneuvers have to be analysed, simulated and tested extensively. Currently DLR is building up a simulation facility to provide test and verification capabilities for rendezvous and docking processes of on-orbit servicing missions. It will be a hardware-in-the-loop simulation facility for physical real-time simulations of rendezvous and docking maneuvers. This test bed will allow the simulation of the last critical phase of the approach process, 25 meters to 0 meters, including the contact dynamics simulation of the docking and berthing processes. The different aspects and challenges of the DEOS mission profile in the area of flight dynamics are presented in this paper. In a second part the hardware-in-the-loop simulation of rendezvous and docking maneuvers is described.
2. THE DEOS MISSION CONCEPT A preliminary mission concept of the DEOS mission has been worked out by SpaceTech GmbH (STI) within a phase-A study [4]. The resulting mission profile is presented in this chapter.
2.1 Overview The DEOS space segment consists of two spacecraft, denoted now as client and servicer. It is planned that both satellites will be injected into an initial polar orbit (600 km altitude) by the same launcher. Fig.1 shows the basic space and ground segment components for the implementation of the DEOS mission.
It is foreseen that GSOC as mission control center together with Weilheim as ground station will build the core elements of the ground segment. They can be augmented by a suitable supplementary ground station network to cover critical mission phases as LEOP (Launch and Early Orbit Phase) or proximity operations and experiments. Even additional communication links via geostationary relay satellites and their operating ground stations are envisaged to improve the performance of the DEOS mission and to enhance operational capabilities. Both client and servicer will have their own communication links to ground (TM/TC), where the servicer is capable to receive the client’s telemetry for on-board autonomous supervision and evaluation purposes.
2.2 Mission analysis
2.2.1 Orbit selection It is currently planned to launch the DEOS servicer and client configuration into an initial polar orbit with an altitude of 600 km and an inclination between 85° and 90°. During the mission, this orbit will be lowered stepwise down to a height of 400 km. Experiments will be performed on all main altitude levels. Depending on the nodal drift, identical illumination conditions repeat every 4.3 to 6 months. The last operations will be the atmospheric re-entry initialized on the 400 km orbit by a corresponding maneuver. According to the functional specification of the DEOS mission [3] it is required that the complexity of the experiments to be conducted in space shall be stepwise increased over the mission period. This means also, that the environmental conditions for the experiments shall be less demanding at the beginning of the operational phase. For example the experiments especially operated with camera based relative navigation techniques including client motion estimation will be conducted first under favorable illumination conditions of the client with a repetition later on under worse conditions. Finally the selection of the initial orbit depends on the duration of the commissioning phase (about 1 month), the selected sequence of experiments and the experiments requiring specific environmental conditions as contact periods or illumination.
2.2.2 Far range formation flying During the period of far range formation flying absolute navigation sensors as GPS or the use of angle tracking data serves as a reference for navigation purposes. Servicer and client are flying in a nominally constant distance from each other within predefined tolerances, where the distance between both spacecraft has to be 2 km as minimum. The status of far range formation flying can be considered as a kind of safe mode functionality with respect to collision risks even without ground control over a couple of days. This allows servicer and client also to stay in this mode for several months when they are not within an experimental phase. Furthermore it is possible to determine the dynamical behaviour of each individual spacecraft and to evaluate the spinning and tumbling motion especially of the client. The end of the far formation flying period represents simultaneously the initialization of the rendezvous operations.
2.2.3 The rendezvous phase Main task of the rendezvous phase is to bring the servicer from far formation flying distance into the close vicinity of the client. Fig.2 shows the scheme of the sub-phases during the rendezvous period. It includes the hold points that are considered as mandatory for the camera based navigation approach to an unknown target satellite.
Far Formation Flight
RV Entry Gate
Phasing Far Range Rendezvous Close Range Rendezvous
Final Approach
Hold Point
Safe Point
Mating Point
Close Hold Point
Fig.2. Sub-phases of the Rendezvous Period
The servicer reduces the distance to the client via several Hohmann-like orbit maneuvers that end at distinct hold points. The final approach to the mating position will be performed via a v-bar maneuver. The location of the v-bar maneuver, either starting from safe point or close hold point, has to be finally determined in the next phase of the DEOS project. The rendezvous phase is split into the following sub-phases: Phasing: Starting from the far formation flying distance the servicer reduces the phase angle to the client. During this phase operations of the servicer are based on absolute navigation techniques. The phasing terminates at the rendezvous entry gate. Far Range Rendezvous: During the far range rendezvous phase the servicer is transferred from the rendezvous entry gate to safe point which is in the vicinity of the client. This is also the starting condition for the close range rendezvous approach. However, if the relative navigation process is based on camera images, at least one intermediate hold point is required at the distance of 75 meters to the client. This allows the controlled switch-over from the far range to the mid range navigation cameras. Close Range Rendezvous: A first target motion estimation is conducted with respect to the client at the safe point followed by a fly around. This is necessary before the docking or berthing procedure with respect to a more or less unknown target can be performed. Afterwards the close range rendezvous process is started with an orbit maneuver from the safe point to the close hold point. Arrived at the close hold point the navigation cameras have to be switched from mid range to close range mode. From the close hold point the final approach to the docking or berthing box is performed via a straight line trajectory (v-bar maneuver). During the final approach the servicer has to be kept within a corresponding pre-defined corridor in an onboard automated way. If the approach corridor is violated a collision avoidance maneuver has to be initiated autonomously onboard. The close range rendezvous terminates at the hold point at the berthing / docking box denoted as mating position.
2.2.4 The capture and berthing process The end-effector of the servicer manipulator arm is grappling the client. The client is then latched onto the berthing port of the servicer. For the DEOS mission the servicer’s berthing port is identical with the servicer’s docking port. The individual steps of the berthing procedure are described in Fig.3.
Acquisition of Berthing Box
Tracking
Grappling
Steering
Latching
Fig.3. Berthing Procedure
Acquisition of Berthing Box: The berthing box is defined by the maximum range of action of the manipulator arm. The envisaged mating position for the berthing act is maintained using the relative navigation sensors. Tracking: The target to be grappled, client respectively, is in the field of view of the manipulator arm cameras. It is either tracked by the ground operator via telematics operations or supervised autonomously onboard by onboard image processing. In this way the final position before grappling can be achieved. Grappling: The end-effector is grappling the target, client respectively, at the dedicated grapple fixture. This can again be conducted by the operator on ground via telematics operations or supervised autonomously onboard. One of the primary mission goals of DEOS is achieved with the successful capture of a non-cooperative client satellite. The following two steps are additional mission goals to include also a cooperative client spacecraft. Steering: After grappling of the client, the client has to be steered to the berthing port for latching. This is performed by the manipulator arm and usually supervised autonomously using the docking camera images of the client’s marker LEDs. Latching: This is the final step in order to achieve a rigidly coupled configuration. The berthing interface of the client is inserted into the servicer’s berthing port, while the drive of the servicer’s berthing mechanism operates actively in a secure way in clamping the clients berthing interface. The latching process is usually performed autonomously.
2.2.5 The docking process for a cooperative client The docking process is based on the reaction control system and the images of the servicer’s docking camera. The client is being approached slowly by the servicer, which is starting from the mating position. The client’s docking interface is inserted into the servicer’s docking port. If the docking is performed via telematics operations from ground also relative position and attitude information of the client is provided by ground processing. In case the docking is done onboard autonomously the relative navigation information is based on onboard processing of the docking camera images. The individual steps of the docking procedure are described in Fig.4.
Acquisition of Docking Axis
Reception
Capture
Latching
Fig.4. Docking Procedure
Acquisition of Docking Axis: The servicer is slowly moving along the docking axis towards the client. The marker LEDs on the client’s docking interface are within the field of view of the servicer’s docking camera. So the relative position and attitude information can be determined either autonomously on board or on ground. Based on the results of the image processing the client’s docking interface is kept properly aligned with the axis of the servicer’s docking port by means of the servicer’s reaction control system. Reception: The reception phase starts when the client’s docking interface enters the cone of the servicer’s docking port. Now it can occur that both docking interfaces contact each other. However, these contacts will result only in small counter rotations of both spacecraft due to the specific shape of the interfaces. No significant repulse forces are expected due to low operating velocities. This is the reason why the approach of the servicer to the client can continue despite these contacts. Capture: As soon as the client’s docking interface is inserted sufficiently into the servicer’s docking port spring driven clamps are pushed aside by the front part of the client’s docking interface. Once the front part has passed the clamps, the clamps lock the front part prohibiting an escape of the client. Latching: The clamps are moved by a drive to obtain a structural connection and a rigidly coupled configuration of two spacecraft.
2.2.6 Flying in coupled configuration and re-entry Basically there are two types of coupled configurations between client and servicer. A rigid coupled configuration can be generated by the docking port or the manipulator mechanism. A dynamically coupled configuration is only possible via the manipulator arm. The latter one will be established for any steering phase during the berthing process and for the determination of dynamical parameters by applying manipulator arm movements. During these periods the active attitude and rate control systems of the servicer and the client are disabled. The rigidly coupled configuration is especially used for the execution of different experiments and de-orbiting maneuvers. Here simulations have been performed in order to demonstrate the feasibility of the attitude control tasks for these configurations within the required limits. The selected sequence of de-orbiting maneuvers transfers the coupled configuration of servicer and client from a circular orbit to a lower point by decreasing its perigee height. servicer and client are linked together both via docking and berthing port where the manipulator arm is providing sufficient contact pressure between the two spacecraft with all manipulator arm joint brakes activated. The client is flying in front of the servicer and is in charge of controlling the re-entry maneuver. The most efficient way to decrease the perigee altitude of an orbit is to perform a braking maneuver at the apogee such that the apsidal line of the orbit remains unchanged. The maneuver can be done at any location on the orbit, if the initial orbit is a circular one. So it is possible to fix
the apsidal line orientation by the maneuver location. This is an essential point since the impact latitude is directly correlated with the apsidal line direction. Starting from an initial circular orbit a set of successive maneuvers performed at the apogee can be used to decrease the perigee altitude. The envisaged longitude will be achieved by initiating the final de-orbiting sequence on the appropriate number of orbit. Fig.5 shows the client/servicer re-entry configuration.
3. VERIFICATION OF THE RENDEZVOUS AND DOCKING PROCESSES ON GROUND
One of the critical phases of an on-orbit servicing mission is to ensure a safe and reliable rendezvous and docking (RvD) process. Especially this phase has to be analyzed, simulated and verified in detail. Classical approaches, e.g. numerical simulations, deliver only limited results. Therefore tests or test facilities have to be defined where the entire RvD process including real flight hardware components of the guidance, navigation and control (GNC) system can be simulated and evaluated under utmost realistic conditions with respect to the space environment.
3.1 Experiences in simulating rendezvous and docking maneuvers DLR has more than two decades experience in the field of simulating RvD processes. The former EPOS facility (European Proximity Operations Simulator) was a test bed jointly developed by ESA and DLR for laboratory simulation of rendezvous maneuvers of spacecraft over the last 12 meters of the rendezvous phase (without docking simulation). It consisted of a six degree of freedom gantry robot for simulating the chaser motion, a three axis servo table for simulating the attitude motion of the target and a sun simulator to generate utmost realistic illumination conditions. The last intensive utilization was the test and verification of the European Automated Transfer Vehicle (ATV) RvD sensors and systems and is shown in Fig.6.
Target Pattern RVS
Fig.6. Rendezvous Simulation with the Former EPOS Facility for ATV Sensor Tests The facility was used for the following projects in the field of GNC test and verification in the past:
• 1991 Comparative sensor tests (MATRA / Saab / MBB) • 1997 -1999 ARP tests (ATV Rendezvous Pre-Development Project) • 1997 - 1999 Planetary Lander Test Facility (Study Phase) • 1998 – 2000 EDISON (Distributed simulation over ATM network) • Since 2001 GNC tests in the frame of the ESA ATV project – sensor tests and closed loop
tests of the complete GNC system • Since 2003 tests for the Japanese HTV project – sensor tests
Because the former EPOS facility was outdated, now DLR has built up a completely new simulation facility to provide test and verification capabilities for RvD processes of on-orbit servicing missions.
3.2 The new EPOS facility After dismantling the former EPOS facility DLR began construction to build up a completely new simulation facility. The new facility is shown in Fig.7. The new DLR RvD simulation facility comprises a hardware-in-the-loop simulator based on two industrial robots for physical real-time simulations of rendezvous and docking maneuvers. This test bed will allow for the simulation of the last critical phase (separation ranging from 25 to 0 meters) of the approach process including the contact dynamics simulation of the docking process. The new simulation facility will be characterized by the following:
• It is a highly accurate test bed. The measurement and positioning performance will be increased by factor 10 compared to the former EPOS facility, i.e., the accuracy is better than 0.5 millimeters in position and better than 0.01° in rotation.
• Dynamical capabilities will allow for high commanding rates and the capability of force and torque measurements.
• The simulation of sunlight illumination conditions as well as the compensation of earth gravity force will both be part of the assembly to generate an utmost realistic simulation of the real rendezvous and docking process.
• The utilization of standard industrial robotics hardware allows a very high flexibility with respect to different application scenarios.
Fig.7. The New EPOS facility The new facility consists of the following components:
• A rail system mounted on the floor to move an industrial robot up to a distance of 25 meters. • An industrial robot (Robot 1) mounted aside the end of the rail system for simulating the 6
degrees of freedom of one spacecraft (e.g. client). • An industrial robot (Robot 2) mounted on the rail system for simulating the 6 degrees of
freedom of the second spacecraft (e.g. servicer). • A PC-based monitoring and control system to monitor and control the RvD simulation on
the facility. In can be divided into three levels: o The local robot control where each axis of the robots can be controlled
separately. o The facility monitoring and control system (FMC) where the entire
facility is controlled in real-time. o The application control system which handles the actual RvD simulation
application. The various components are depicted here in Fig.8.
PC-based real-time facility control system
Robot 1 with 6 DOF
• Carrying client satellite mock-up
• Motion simulation of client satellite
Robot 2 with 6 DOF on a 25 m rail system
• Carrying RV sensors and docking system of servicing satellite
• Motion simulation of servicing satellite
Fig.8. Components of the New EPOS-facility
3.3 A typical RvD hardware-in-the-loop scenario A typical set-up of the EPOS facility for a DEOS RvD simulation scenario is shown in Fig.9. In this individual case the client is mounted on the Robot 2 (rail). It must be emphasised here, that the facility is basically designed for both options, i.e. mounting of client and servicer can be switched according to the specific needs of the tests.
Fig.9.EPOS Simulation Set-up for DEOS
For such a hardware-in-the-loop simulation scenario the RvD sensors and the robotic manipulator arm are mounted on one robot and a typical satellite mock-up of the client satellite is mounted on the other robot. The RvD sensors can measure the relative position and attitude of the client satellite. Based on that information the onboard computer calculates the necessary thrusters or reaction wheel commands. These will feed into a real-time simulator. The dynamics simulator computes then an update of the state vector (position and attitude of the two spacecraft) for the next sample based on all relevant environmental and control forces, torques respectively. Then the state vector for the new sample will be commanded to the facility. This represents a characteristic scenario and is depicted in Fig.10.
Robot 1
Robot 2
Measurement System
(e.g. Camera)
Relative position and attitude measurements
Position/attitude control commands, e.g. thruster commands Application Control
3.4 Project status and future planning After dismantling the former EPOS facility in 2008 the design of the new facility was started. The new robotic hardware was installed in January 2009. In parallel the development and implementation of the entire facility monitoring and control system was started and will be finalized in September 2009. At the end of this period a first baseline concept of the RvD simulator is available. Afterwards the following upgrades are planned:
• Implementation of an Online-Measurement-System to achieve higher performances in accuracy.
• Implementation of the contact dynamics simulation capability with respect to the docking process.
• Implementation of a sun simulator for utmost realistic simulation of illumination conditions.
4. CONCLUSION The present paper reflects the results of the phase-A study of the DEOS mission which has shown the feasibility of the mission. With the EPOS facility DLR/GSOC is able to provide the means for verification and testing of the servicer’s docking-payload. A more detailed analysis in the area of flight dynamics will be performed in the upcoming phase-B and will be described in future presentations.
5. ACKNOWLEDGEMENT The DEOS mission phase-A study has been awarded by the Space Agency of the German Aerospace Center (DLR) and funded by the German Federal Ministry for Economics and Technology (Förderkennzeichen 50 RA 0802).
6. REFERENCES [1] Boge T., et al., Hardware-in-the-loop Simulator für Rendezvous und Docking Manöver, DGLR,
Deutscher Luft- und Raumfahrtkongress, Aachen, Germany, Sept. 2009 [2] Boge T., Schreutelkamp E., A New Command and Control Environment for Rendezvous and
Docking Simulations at the EPOS-Facility, 7th International Workshop on Simulation for European Space Programmes - SESP 2002, Noordwijk, Netherlands, November 2002,
[3] DEOS Functional Specification, DLR / Space Tech GmbH (STI), January 2009 [4] DEOS Mission Description Document, Space Tech GmbH (STI), January 2009 [5] Fehse, W., Automated Rendezvous and Docking of Spacecraft, Cambridge University Press,
2003 [6] Rupp T. and Boge T., Introduction of the new RvD HIL Simulation Facility ARDOS, 3rd
International Workshop on Verification and Testing of Space Systems, Torino, Italy, April 2009
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