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Further options for GRACE attitude data processing

Improved star camera attitude data 5

5.6 Further options for GRACE attitude data processing

The GRACE SCA1B RL02 data contain systematically higher noise than expected due to the imperfect implementation of the method for SCA data combination in the JPL processing routines, as we have proven in this chapter. As stated by Kruizinga et al. (2013), the correct implementation of the combination method will be a subject for the the next data reprocessing, if any. Further improvement of the GRACE attitude data, however, is still possible. It is based on further refinement of the star camera data processing and also on the attitude data fusion from multiple sensors. In the following, a brief discussion of the individual processing steps is presented. The implementation of these steps in the real data processing is, however, beyond the scope of this thesis.

In the currently implemented method for SCA data combination it is assumed that the measurement accuracy of the two star cameras is the same. But as we showed in Section 3.3.4, this is not the case. Therefore, by taking these performance differences into account, further improvement of the combined solution could be obtained. In the combination method based on the weight matrix (cf. Section 5.2.1) setting of the parameters , κ, Pindividually for each SCA head is possible. The weight matrix P given in Eqaution 5.3 will then be adjusted to:

Pi =


2x 0 0

0 1

2y 0

0 0 1



Some of the key parameters for the attitude data processing and for the derivation of the inter-satellite pointing angles are the QSA, QKS and VKB calibration parameters. It is most likely, that these parameters were estimated based on the non-optimally combined attitude quaternions. Therefore, using the improved SCA attitude data, the reprocessing of the observations from the KBR calibration maneuver could lead to better results.

Another feature related to the star camera data are jumps in the SCA1B data at transitions from the combined SCA solution to single camera solution and vice versa, see Figure 5.16. This figure shows the inter-satellite pointing pitch angle, which was derived from both SCA head#1 and #2 data sets and from the combined SCA attitude solution. The jumps at transitions from single to dual camera data reach up to 0.3 mrad and are the consequence of the different performance of the two star camera heads. These jumps remain to be untreated in both SCA1B RL01 and RL02 data. They could be reduced by applying e.g. a smoothing filter.

Figure 5.16: Jumps in SCA1B data at transitions from dual to single camera mode. The figure shows the pointing pitch angle computed based on SCA1A head#1 data (red), SCA1A head#2 data (blue) and SCA1B

data (black). cUng-Dai Ko, CSR

The information about the spacecraft’s attitude is provided not only by the star cameras, but also by the accelerometer and the IMU, which measure the satellite’s angular accelerations and angular rates, respectively. Both of these sensors are characterized by lower measurement noise in the high-frequency band that the SCA, the long-term stability of the data is affected by drifts and biases, though. Therefore a frequency dependent attitude data fusion is possible and would further increase the attitude accuracy.

The combination of the star camera and accelerometer data was already tested by Frommknecht (2008). The data were combined on the level of angular rates by means of low-pass filter for

the star camera data and high-pass filter for the accelerometer data with a cut-off frequency of 3·10−2Hz. Similar method, using Wiener filtering, was implemented for the combination of GOCE star camera and gradiometer data (Stummeret al., 2011). Recently, the GRACE SCA/ACC data combination was implemented by Klinger and Mayer-Gürr (2014). The au-thors obtained the combined attitude solution not by data filtering but as a result of variance component estimation. Theoretically, the combination of the star camera data with the IMU data would be also possible. However, the IMU on GRACE-A failed right after launch and the IMU on GRACE-B is turned off most of the time. Therefore only the fusion of GRACE star camera and accelerometer data is practicable.

that makes a big difference.

Sir Winston Churchill

-Attitude determination and mission lifetime 6

The GRACE mission lifetime is limited by several factors which are discussed in detail by Herman et al. (2012). One of the factors limiting the mission lifetime is the amount of propellant (gaseous nitrogen GN2) onboard the satellites. In the absence of propellant, the inter-satellite pointing cannot be maintained and thus the inter-satellite ranging observations needed for the gravity field recovery cannot be collected. Also the communication with ground stations would be impossible if the RX/TX antennas would not be orientated towards the Earth as needed. The mission lifetime further depends on the energy which keeps the satellites’

instruments, sensors and computers alive and which allows the performance of all measurements and communication with ground stations. The energy budget depends on the capacity of batteries and performance of solar cells. Another limiting factor is the thruster operation, which is guaranteed by the manufacturer up to 106 activations cycles. As the GRACE orbit is designed to be freely decaying, the satellites’ altitude affects the mission lifetime as well. Lower orbit means higher air drag and disturbance torques due to the residual atmosphere, which would critically affect the mission performance. Also, the degradation of the KBR assembly due to front-end oxidation is another factor influencing the observation period. Generally, the mission operation depends on the health of the whole onboard laboratory.

It is a great accomplishment of the GRACE mission operations team to have kept the GRACE mission nominally operating despite several defunct components for more than 8 years after the planned mission termination. The very good performance of the twins is a result of continuous optimization, parameter adjustment, adaptations, software update, satellite maneuvers, etc. The current greatest challenges are to keep the energy budget stable and to optimize the propellant consumption. Demanding spacecraft maneuvers and handling are necessary to optimize the battery performance after 2 solar cells failed on each spacecraft and the battery capacity decreased from nominal 16 Ah to 3 Ah (Herman et al., 2012). In order to minimize the propellant consumption, several approaches have been tested and implemented on GRACE, cf. Section 6.3. Also the health of the scientific instruments such as the K-band ranging assembly, accelerometer, GPS receiver and the star cameras is under critical observation.

Although the nominal limit of 106 thruster activation cycles has been exceeded by some of the 12 attitude control thrusters (see Table 3.2), so far they continue to work nominally.

The relation between the mission lifetime and the attitude determination may not be very obvious at the first sight. However, the accuracy of the in-flight determined satellite’s attitude critically affects the propellant consumption and the number of thruster activation cycles needed to keep the satellites in their required orientation. In this chapter, we present the impact of the different performance of the two star camera heads on both the propellant consumption and the number of thruster activation cycles, which are both considered as important factors limiting the mission lifetime. As first, we discuss the accuracy of the inter-satellite pointing angles, which were determined based on single star camera data (cf. Section 6.1). This is followed by demonstration of how the propellant consumption and thruster operation depends

on the primary star camera (see Section 6.2). Finally, options for reduction of the propellant consumption are presented in Section 6.3.

6.1 Accuracy of inter-satellite pointing angles derived from single