DEVELOPMENT OF AN INFRA-RED BREAK UP RE-ENTRY SENSOR

Transcription

1 DEVELOPMENT OF AN INFRA-RED BREAK UP RE-ENTRY SENSOR R. Greger (1), Chr. Benkeser (2), W. Coppoolse (3), Th. Roesgen (4), N.P. Murray (5),H.Weihs (6), Chr. Dittert (7), M. Vigano (8), N. Ortiz (9) (1) RUAG Space, Schaffhauserstrasse 580, 8052 Zurich, Switzerland, (2) RUAG Space, Schaffhauserstrasse 580, 8052 Zurich, Switzerland, (3) RUAG Space, Schaffhauserstrasse 580, 8052 Zurich, Switzerland, (4) ETH Zurich, Sonneggstrasse 3, 8092 Zurich, Switzerland, (5) ESA ESTEC, Keplerlaan 1,2201 AZ Nordwijk ZH, The Netherlands, (6) DLR, Pfaffenwaldring 38-40, Stuttgart, Germany, (7) DLR, Pfaffenwaldring 38-40, Stuttgart, Germany, (8) ViaSatAntenna Systems SA, EPFL Innovation Park Bât J, 1015 Lausanne, Switzerland, (9) ViaSatAntenna Systems SA, EPFL Innovation Park Bât J, 1015 Lausanne, Switzerland, ABSTRACT During the re-entry of the Automated Transfer Vehicle (ATV) the break-up and heating up of the ATV structure shall be monitored to gain information and to improve the theoretical models, used to predict the reentry of the International Space Station (ISS). Gathering in-situ data, however, is highly challenging mainly due to the harsh environmental conditions during structural break-up of the spacecraft impacting also the integrity of data recording and data downlink devices. For this task a modular, autonomously operating infra-red camera system prototype has been developed and built within only 9 months after kick-off. The camera system, mounted to the ATV cargo rack system, is activated by the ISS Crew before ATV de-docking and autonomously detects ATV s final re-entry accounting for mission events such as, orbit correction boosts or mission delays without running low of power. 1. INTRODUCTION Re-entry predictions for large space vehicles are usually based on theoretical models. Predictions of the structure breakup and the resulting debris field are still uncertain. This is due to the fact that gathering experimental data in the harsh environment during re-entry needed to prove and develop the theoretical models, is very challenging. Within the clean space initiative, attempts are made to improve theoretical models to develop new space craft designs to ensure a complete breakup of the structure keeping the impact on the environment low (Design for Demise). Furthermore re-entry scenarios for the ISS, the largest man-made structure in space, have to be developed and optimized to reduce the impact on Earth to the maximum extent. Figure 1: Artist View of ATV-5 Re-Entry (Credit ESA) The re-entry of the last Automated Transfer Vehicle (ATV) George Lemâitre, offered the unique possibility to test a new re-entry trajectory more comparable to the latter ISS re-entry trajectory and to monitor the heating up and final break-up of the ATV structure using in situ data acquisition. In May 2013, ESA decided to join the last ATV campaign with the first new developed European re-entry recorder, the ATV Breakup Camera (ATV-BUC). Figure 2: ATV-Cut Away (Credit ESA) A Re-Entry recorder is designed to operate as a type of black box, acquiring sensor data during the re-entry

2 phase and transmitting the data to ground via available communication networks until it is destroyed during splash down in the ocean. For the final ATV-5 re-entry campaign, three complementary re-entry recorders installed in ATV s Cargo Bay prior to ATV s de-docking were foreseen. REBR, an US recorder acquiring sensor readings, I- Ball, a Japanese recorder capturing images in the visible frequency range, and ATV-BUC, the European recorder capturing video images in the Near Infra-Red (NIR) frequency range of ATV s front hatch heating up until breakup of the Cargo Bay s hull. Unfortunately I-Ball was lost in the Cygnus launch disaster and could not be recovered in time. Being the maiden flight of the European re-entry recorder the main objective of the ATV-BUC was to prove the overall generic recorder design concept by transmitting any captured data recordings. The final ATV5 re-entry campaign offered the unique opportunity to capture an IR video sequence to complement the scientific re-entry data captured by REBR and I-Ball. The ATV-BUC comprises two modules: The Infra-Red Camera (IRC) front-end capturing the temperature distribution of the cone and hatch area until the camera will be destroyed during breakup The SatCom unit receiving the data i.e. images from the front-end and transmitting the data via the Iridium network during the final re-entry phase. It is designed to survive the ATV breakup, but it is not expected to survive the final splash down. Figure 3: IRC Field of View within ATV Before ATV leaves ISS the camera system is mounted to the ATV cargo racks in order to cover parts of cone and hatch within its Field of View (FOV) of 30 as shown in Fig.3. Before closing the hatch the ISS Crew activates the equipment. The ATV-BUC runs autonomously during the re-entry mission using its internal battery pack. This implies that the activation sequence has to cope with mission events like orbit correction boosts or mission delays due to unforeseen events without running low in power. The ATV5 flight schedule imposed the biggest challenge on the development schedule: Kick-Off: 15. July 2013 Concept Review: 15. August 2013 Design Review: 20. September 2013 Acceptane Review: 07. March 2014 Delivery: 14. March 2014 Launch: 29. July 2014 RUAG Space as prime of the industrial consortium was responsible for the design, testing and verification of the IRC and the SatCom. DLR developed and provided the Thermal Protection System (TPS) for the SatCom. ViaSat designed the Iridium antenna system and the SatCom electronics housing. GOMSPACE delivered and qualified the on-board batteries and ETH Zurich developed the IR front-end and the SatCom software. 2. DESIGN CONCEPT One objective of the ATV-Break-Up camera is the capturing of a video of the heating up of ATV s structure using an Infra-Red camera system during reentry, it is required to function until the ATV structure breaks. The camera system is based on the camera system also flown on ESA s IXV mission. Due to the fact that ATV will be destroyed during re-entry the captured image data has to be downloaded via an ATV independent downlink during the re-entry phase. Consequently the ATV-BUC is equipped with a re-entry body (the SatCom unit) surviving the hot phase of the re-entry. It provides intermediate storage of scientific data and transmitting capabilities to ground until its final destruction. The downlink of choice is the Iridium network which provides a worldwide downlink coverage and which has been used by the REBR in previous missions. The ATV-BUC concept is shown in Fig.4.The IRC unit manages the experiment switch-on sequence dependent on ATV s mission time line, captures IR images of ATV s forward cone and hatch, stores the raw image data and forwards it to the SatCom unit. The SatCom unit itself is designed as a spherical reentry body to withstand the high thermal loads and the harsh environment during re-entry. Its task is to receive and store the scientific data (in this case images), to establish the downlink via the Iridium network, to compress the raw video stream and to transmit the compressed data until the HW is destroyed or the battery is down. The ATV-BUC is transported to ISS as payload and installed and activated by the Crew just before ATV s de-docking for re-entry. The ATV-BUC enters a

3 hibernating mode until the final orbit correction boost of ATV for the final re-entry is detected. Main Switch On Timer (Hibernating) IRC SatCom Re-Entry Detector (Thermo-Switch) Temp Monitoring Accelerometer IMU Break_Up Detector IR Camera Accelero meter Timer Main Switch Battery, 14V IRC DHU Pyrometer Switch_On_Logic Test Interface Ethernet Power On 12V_Ext Test_IF Switch_Off CPU & Storage Switch On Logic DC/DC External Power Switch Figure 4: BUC Concept SatCom Antenna Transceiver Transceiver Switch On Logic Battery, 14V Power On Break-Up Detector (Baroswitches) The IRC Handling Unit (DHU) sets up the IR camera head after booting and starts capturing images of ATV s cone and hatch. The standard frame rate is set to 10 frames/sec to allow exposure time corrections during heating up of the target structure. The images are forwarded to the SatCom unit and stored until the downlink is established. The IRC is also equipped with a contactless temperature sensor providing additional temperature readings to the DHU. This temperature sensor acts also as a back-up camera trigger to power-up the IRC when the structure starts to heat up. The heart of the SatCom unit is a Single Board Computer (SBC) connected to the Iridium modem. After power-on of the SatCom the SBC connects to the IRC DHU and transfers the captured data into its own memory. After interruption of the video stream (no more images received) it activates the modem and tries to connect to the Iridium Network using Iridium s Short Burst (SBD) service. Simultaneously the video stream compressor is started. If the link is available the compressed image data is merged with auxiliary TM data, packetized into SBD messages and forwarded to the modem for downlink. In the functional diagram Fig. 5, arrows in green colour indicate the power/switch-on flow, in blue colour a state transition and in red colour the data flow. Fig. 6 shows the ATV-BUC flight unit. 3. MISSION SZENARIO/TRIGGER CONCEPT The ATV-BUC shall monitor the heat-up of the ATV structure, i.e., forward cone and forward hatch as part of the pressurized Cargo Bay until break-up of the structure. The acquired data has to be transmitted to the ground. During break-up of the HW the camera system is most probably destroyed. Consequently the acquired data has to be stored in a module which survives the ATV break-up and which re-enters into the atmosphere and transmits the scientific data to the ground. Fig.7 provides an overview on the ATV-BUC mission scenario. IR-Camera CMD Switch on Boot age Capture Image Buffering Until SatCom boot finished Transfer «rsync» Boot Switch on Receive & Store Compress Transfer Re-Entry Detection Modem Control Start Switch On Transceiver Establish DHU SCC Transceiver Figure 5: Functional Diagram Figure 6: ATV-BUC Flight Unit Based on this scenario the system has to fulfil the following main requirements: Detection of the tumbling phase of the ATV in which the structure starts heating-up Power up of camera when final tumbling phase is detected Storage of scientific data until transmission to ground is possible No transmission or radio frequency emission before final tumbling phase (ATV requirement, 3 independent inhibits) Detach SatCom form ATV during re-entry Ensure that the batteries provide sufficient power to run the equipment until the end of the mission, i.e., most of the data is transmitted to the ground. Ensure the correct operation of SatCom unit during break-up/re-entry although the IRC front-end may not work anymore Ack Modem Setup Link available

4 Activation & De-Docking Final De-Orbiting Boost Orbiting Phase (<15 days) ATV Heating-Up ATV Break-Up Splash Down seconds in both cases. The heat-up of ATV is expected to start at an altitude of 120 km until the final break-up Hibernation & Waiting for Final Boost IR-Camera & SatCom Activation IR-Video Capture & Transfer to SatCom Altitude [km] Nominal Shallow Shallow - shift End of IR-Camera SatCom Ejection & Iridium Connection Trials Time [sec] Transfer to Ground via Iridium Splash Down & End of Mission Figure 7: ATV-BUC Mission Scenario During the ATV-BUC development phase the mission timeline for the ATV-5 re-entry has been kept open. The decision to change the standard re-entry trajectory (Nominal) to a trajectory as proposed for the ISS (Shallow) was taken a few weeks prior to the planned departure of the ATV from the ISS. Fig. 9 shows a comparison between Shallow and standard trajectory. For the Shallow trajectory, a re-entry campaign similar to ATV-1 to observe the re-entry by ISS and airplanes is set up. Figure 9 Possible ATV Re-Entry Trajectories at an altitude of 75 km as shown in Fig. 10. The expected capturing time for the images is then 600 seconds for the shallow and 500 seconds for the standard re-entry trajectory. The SatCom unit itself has to start the downlink at an altitude of approximately 75 km when the ATV Cargo Bay hull breeches. However, a stable Iridium link to download the data is only guaranteed for a maximum SatCom speed of less than Mach ATV Breakup Altitude [km] Start Nominal Shallow Shallow - shift Time [sec] Figure 8: Re-Entry Campaign Logo This campaign constrains the re-entry time-line by the fact that the re-entry area has to be located in the FoV of the ISS and simultaneously in the operational range of the observation aeroplanes. A detailed prediction of the mission time is therefore only possible just before the re-entry is commanded. Consequently the camera trigger algorithm has to be able to cope with time shifts up to 15 days operational time. Note that in the following figures a Shallow-Shifted curve has been added that shows the Shallow trajectory shifted in time, in order to simplify the comparison between the two possible trajectories Shallow and Nominal. The real re-entry phase is lasting approximately 3000 Figure 10: ATV-Re-Entry Detail The trigger concept is based on the use of different trigger sources to provide redundancy and synchronizing with the ATV mission timeline using the orbit correction boosts as references. After activation of the ATV-BUC it enters a timer controlled hibernation mode in which a low power accelerometer and a pyrometer looking towards the ATV cone are powered. The pyrometer alarm output is pre-set to a temperature threshold well above the ambient Cargo Bay temperatures (e.g., 50 C) to avoid any pre-triggering. Accelerometer and pyrometer trigger are logically or -ed to generate the main trigger for the DHU to provide redundant trigger sources.

5 For the final re-entry de-orbiting boost a fixed boost duration is defined which allows for an unambiguous identification among other orbit correction boosts, which have the same acceleration amplitude (0.07m/sec 2 ) but different boost durations. ATV Launch & Mission Installation in ATV Activation ATV s region of interest is monitored with a Commercial Off The Shelf (COTS) contactless pyrometer. Tab. 2 provides an overview on the most important technical parameters of the pyrometer Table 1: IR camera parameters. Parameter Value Camera XEVA XS Sensor Technology InGaAs Photodiode Array, uncooled Spatial Resolution 320 x 256 pixel Wavelength Range NIR ( µm) De-Docking Hibernation Table 2: Pyrometer Parameters Orbit Boosts Parameter Value Waiting for Final Boost Temperature range: -40 C to 400 C Final Boost ATV Heating Up Start-Up IRC & SatCom Spectral range 8 to 14 µm Accuracy Resolution ± 1.5 C 0.1 C ATV Break-Up Capture End Of IRC Opening Angle: 1.6 C Release SatCom Start-Up Modem Trial to Establish The pyrometer is mounted directly below the IR camera to ensure that the measuring spot of the pyrometer coincides with the FoV of the IR camera (see Fig. 12) Blackout Establish Splash Down End Of SatCom Figure 11: ATV-BUC Mission Flow Chart The SatCom and the IRC will be activated simultaneously to provide a protected data storage inside the SatCom. The ATV safety rules require that the Iridium modem cannot be powered until ATV enters the tumbling phase. To comply to these rules the Iridium modem is inhibited by several independent switches (thermal, pressure, acceleration) in the modem s power lines, which allow booting the modem only if the SatCom outer sphere is heated up due to its own re-entry. 4. IR-CAMERA The IRC module uses an uncooled, low weight, near Infra-Red camera (XEVA XS) with low power consumption as camera front-end (see Tab. 1). It features variable exposure times ranging from 4µs up to 200ms, allowing to measure an extended range of temperatures from 500K to 2000K. The camera is operated by a separate Handling Unit (DHU) based on a design which has been successfully flown on the IXV. The temperature of the Figure 12: IR-Camera and Pyrometer Assembly The IRC module is powered by a Li-Ion based battery developed and qualified by Gomspace (DK) for CubeSat applications. The battery and the ATV-BUC mission controller are integrated in a separate housing as shown in Fig. 13 Figure 13: IRC Module(Left: Battery, Right: DHU) 5. SATCOM UNIT The SatCom unit comprises a spherical re-entry body and holding bracket for accommodation inside ATV s

6 Cargo Bay. Fig. 13 shows the SatCom module after testing. Figure 16: Antenna Radiation Pattern. Figure 14: SatCom Flight Unit The SatCom is controlled by a COTS Single Board Computer (SBC) which interfaces to an Inertial Measurement Unit (IMU), the Iridium modem, the power supply and the outer world via a standard Ethernet link. The modem output signal is splitted and routed to two patch antennas at the poles of an 180mm diameter Aluminium sphere hosting the electronics. The IMU provides information on accelerations, rotation rates and magnetic field to the SBC. The power supply based on a Li-Ion battery is designed for a minimum operational time of 30 minutes. Fig. 15 shows the electronics accommodation inside the sphere. For the re-entry phase an omnidirectional radiation pattern of the communication antennas is favourable because the SatCom will not have a preferred flight orientation. The achieved radiation pattern and gain analysis for the final antenna system configuration shows a nearly omnidirectional radiation pattern except the equatorial plane (see Fig. 16). Figure 17: Simulated Surface Temperature of the Whipox Sphere During Re-Entry Starting at an Altitude of 85km The temperature of the Whipox sphere s surface increases fast to a maximum value of 1800 C. At the end of the mission it will be approximately 580 C (see Fig. 17. The performance of the insulation is proven looking at the maximum temperature at the inner boundary of the insulation which does not exceed 38 C as shown in Fig. 18. After an operation time of about 665 seconds the temperature of the inner sphere reaches 38 C taking into account also the self-heating due to the internal electronics power dissipation which is well within the limit of 80 C. Figure 15: SatCom Inner Electronic Sphere For the re-entry the inner sphere needs protection against the occuring heat fluxes and plasma flows. For this purpose the inner sphere is encapsulated in a thermal protection system comprising a Whipox (Wound HIghly Porous OXide) ceramic sphere (30mm diameter) and high performance ALTRA MAT 72 insulation. The thermal analysis of the SatCom sphere is based on the heat fluxes impacting the sphere for the standard and shallow re-entry as well as the heat dissipated by the electronics inside the inner sphere. Figure 18: Thermal Analysis Results Fig. 19 shows images of the SatCom Flight HW during assembly. On the left side the closed SatCom electronics sphere, on the right the SatCom unit before final closure of the Whipox sphere. Both Whipox half spheres are connected with C/C-SiC rivets. Figure 19: SatCom Sphere Before Closure

7 The SatCom sphere is fixed with two straps to the hinges fixed to an Aluminium baseplate with rotation joints to support the release of the sphere during reentry. When the temperature of the SatCom environment rises the straps will melt, the hinges open and the sphere is released. 6. IMAGE PROCESSING The IRC front end DHU (F-DHU) controls the IR camera and adjusts the exposure times for optimal recording conditions. The images are pre-processed (background subtraction) to reduce the fixed pattern noise prior to further processing. camera can be derived which includes the known sensor spectral sensitivity and bandwidth. This provides a mapping function between temperature ratios and intensity ratios. the true temperature of the reference pixel is determined based on the information available from the IR pyrometer readings because it is pointing at a part of the ATV hull visible also to the IR camera. Additionally it acquires additional sensor data, i.e. IR pyrometer and internal temperatures, and prepares the data for the downlink. The raw image data are sent via the module interconnecting the data link to the DHU (D-DHU) inside the ceramic sphere. This DHU disassembles the raw data stream, performs the image compression and transmits the compressed data in packetized form to the ground via the satellite modem, after establishing contact with Iridium. The image frames are fed sequentially into the high rate image compressor (H.264 format) which converts the individual data frames into a binary stream of compressed data. Sensor / housekeeping data of the F- DHU are treated separately. They are not compressed but merged with additional local sensor data acquired by the D-DHU, i.e. acceleration, rotation rates, magnetic field, temperature. The compressed image data (output of the x264 compressor) and the augmented sensor data are finally repackaged into a lightweight downlink transmission data frame structure. This packetized structure is designed to allow for easy reception and decommutation of the data on ground. Each packet has a length of 1960 bytes, corresponding to the Iridium modem s capacity for one SBD message. The measured temperature distribution is reconstructed based on the transmitted image information. This is necessarily a process with limited accuracy, because the image compression will most likely have reduced the overall pixel data precision. The temperature can be calibrated using a simple three step process: the raw image data is converted to an image containing relative intensity ratios by comparing all pixel values to a specific reference pixel. Such a pixel ratio computation is possible because the characteristic offset values of each pixel were already subtracted on the fly by the F-DHU. a Planck radiation distribution model for the Figure 20: SatCom Processing Flow The procedure outlined ignores the effect of (possibly changing) surface emissivity on the IR signature recorded by the camera. To reduce this unwanted effect a number of special targets with known, fixed (high) emissivity were placed into the camera s field of view. Assuming that the targets have the same temperature as their surroundings, they can be used to estimate the local emissivity ratios 7. MISSION AND FIRST RESULTS The ATV-5 re-entry schedule had been tailored for the Shallow re-entry trajectory: Installation of BUC 09. February 2015 Activation of BUC 13. February 2015 ATV-De-Docking 14. February 2015 Planned Re-Entry 27. February 2015 Final Re-Entry 15. February 2015 Unfortunately a technical problem with ATV s power bus was detected which could not be recovered. To minimise the risk of losing the ATV during the two week cruising time until final re-entry, it was decided to perform a standard re-entry following the Nominal trajectory 30 hours after de-docking. The final ATV-5 re-entry boost lasted 23 minutes and 20 seconds. At 18:04 (GMT) the ATV spacecraft structure broke-up (Fig. 22) releasing the SatCom unit and splashed into the south pacific ocean. After the SatCom had detected the ATV break-up and powered the modem it starts to communicate with the Iridium network trying to send the following message sequence: Ping (Housekeeping ) Video 1 (Video Reference Frame)

8 TM- (Temperatures, Exposure Settings etc.) Video 2+n The SatCom succeed the first connection to Iridium at 18:08:18 (GMT) by sending out a Ping with its housekeeping data as shown in Tab. 3.These data show that the trigger detected the final de-orbiting boost successfully and the camera has captured a video sequence of ATV s hatch for nearly 10 minutes (5967 framerate). The SatCom was spinning and decelerated when the message was generated. The inhibits of the modem worked as expected and the SatCom needed 4 trials in 54 seconds to get a connection to Iridium. The TPS work perfectly, because the CPU temperature was still at 32 C which is in good agreement with the thermal prediction. Following the simulated trajectory of the SatCom the connection to Iridium was established at an altitude of between 30-35km and a speed of around Mach 1. Acc X [g] Acc Y [g] Acc Z [g] Bx [micro T] By [micro T] Bz [micro T] The following connection to Iridium was lost and the second message including the first video packet could not be transmitted. Further investigations to determine why the transmission stopped and a reconnection was not successful are ongoing. 8. CONCLUSION The first European re-entry recorder to capture in-situ sensor data during the re-entry and break-up of a spacecraft has been developed within only nine months from the kick-off. The re-entry recorder has been combined with an IR camera front-end, the ATV-BUC, to monitor the break-up of the last ATV during its reentry. The design concept of the re-entry recorder proved to be well suited to withstand the harsh environment during break-up and re-entry. All systems worked as expected, however, for future missions a telemetry backup system should be considered to improve the reliability of the communications. Figure 21: ATV_BUC installed in ATV (Credit ESA/NASA) Figure 22: ATV Break-Up seen from ISS (Credit ESA/NASA) Table 3: SatCom Ping Ping No: 4 RunTime [sec] 1282 CPU Temp [ C] 32 Transferred Images: 5967 No of magnetic Field Samples 54 Gyro X [ /sec] Gyro Y [ /sec] Gyro Z [ /sec] ACKNOWLEDGEMENTS A big Thank you goes to the highly motivated team at DLR, ETH, VIASAT and Gomspace and especially the ESA team for the good cooperation and the high responsiveness which was the basis of this successful and very fast development. The project was funded by ESA under contracts /13/NL/PA, /13/NL/PA and /14/NL/PA. 10. ABBREVIATIONS ATV BUC COTS DHU DLR ESA ETH FoV GMT IMU IR/IRC ISS IXV NIR REBR SBC SBD TM TPS Whipox Automated Transfer Vehicle Break-up Camera Commercial off the shelf Handling Unit German Aerospace Center European Space Agency Eidgenössische Technische Hochschule Field of View Greenwhich Mean Time Inertial Measurement Unit Infra-Red/Infra-Red Camera International Space Station Intermediate Experimental Vehicle Near Infra-Red Re-Entry Break-Up Recorder Single Board Computer Short Burst Telemetry Thermal Protection System Wound HIghly Porous OXide