Front Side view of CHAMP with location of instruments
Rear Side view of CHAMP with location of instruments
The following figures show the physical layout of the CHAMP spacecraft with the location of some instruments and subsystems.
Front Side view of CHAMP with location of instruments
Rear Side view of CHAMP with location of instruments
The STAR accelerometer sensor is provided by the Centre National d'Etudes Spatiales (CNES) and was manufactured by the Office National d'Etudes et de Recherches Aerospatials (ONERA). It serves for measuring the non-gravitational accelerations such as air drag, Earth albedo and solar radiation acting on the CHAMP satellite. The STAR accelerometer uses the basic principle of an electrostatic micro-accelerometer: a proof-mass is floating freely inside a cage supported by an electrostatic suspension. The cavity walls are equipped with electrodes thus controlling the motion (both translation and rotation) of the test body by electrostatic forces and thus supports the recovery of the orbit from GPS data and by this the gravity field estimation. By applying a closed loop-control inside the sensor unit it is intended to keep the proof-mass motionless in the center of the cage. The detected acceleration is proportional to the forces needed to fulfill this task.
The STAR accelerometer sensor on its base plate
The sensor unit consists of the accelerometer cage containing the proof-mass and the control electronics as power supplies and servo-loops for the electrostatic forces acting on the test body. The walls of the cage (which carry the control electrodes) consist of Ultra-Low-Expansion ceramics (ULE) to keep the influence of thermal variations on the cage dimensions as low as possible. Six servo-channels (3 for linear and 3 for angular accelerations) act separately along the axes of the accelerometer. The pair of electrodes corresponding to each servo-loop is used both for capacitive position sensing and electrostatic force generation. From the measurement of the capacitive a Proportional-Integrative-Derivative (PID) controller determines the drive voltage to be applied to the opposite electrodes for the generation of the elctrostatic field that stabilises (naturally unstable) loop.
A sine wave detection voltage VD is applied to the proof-mass together with a biasing voltage VP (in order to linearise the acting electrostatic forces) by means of a thin gold wire. Both electrodes attract the proof-mass with Forces F1 and F2 proportional to the gradient of the capacitance and to the square of the electric potential differences between the mass and the electrodes. The output network provides the measurement of the actually applied voltage V which constitutes the analogue output of the accelerometer. This analogue output is digitised by 24-bit Analogue-Digital-Converters (ADC's) and reported together with the readout of the capacitive position detectors, the biasing voltage VP and housekeeping parameters in telemetry. The digitising of the analogue signals, provision of secondary voltages and control of the accelerometer is performed by an Interface and Control Unit (ICU) which interfaces the STAR with the CHAMP spacecraft.
The STAR accelerometer sensor axes are defined via the surface normals
of the parallelepipedic proof-mass as follows:
| X: less sensitive axis, parallel to S/C -z-axis | psi: rotation about x-axis of accelerometer |
| Y: high sensitive axis, parallel to S/C x-axis | theta: rotation about y-axis of accelerometer |
| Z: high sensitive axis, parallel to S/C -y-axis | kappa: rotation about z-axis of accelerometer |
The origin of the measurement frame is the centre of the proof-mass. The orthogonality of the axes of this frame is better than 2.5 10-5 rad.
| Measurement bandwidth | 10-4 ... 10-1 Hz |
| Linear accelerations: | |
| Measurement range | ± 10-4 ms-2 |
| Resolution 1) | < 3 × 10-9 ms-2 (y- and z-axis) |
| < 3 × 10-8 ms-2 (x-axis) | |
| Angular accelerations: | |
| Resolution 1) | 1 × 10-7 rad × s-2 (rotation about x-axis) |
| 5 × 10-7 rad × s-2 (rotation about y- and z-axis) | |
| Temperature stability of bias and scale factor: | |
| Max. temperature coefficient of bias | 5 × 10-8 ms-2 °C-1 (x-axis) |
| 1 × 10-8 ms-2 °C-1 (y- and z-axis) | |
| Max. temp. coefficient of scale factor 2) | 2 × 10-3 °C-1 (x-axis) |
| 5 × 10-3 °C-1 (y- and z-axis) |
1) within the specified measurement bandwidth
2) assuming a maximum temperature variation of 1°C per
orbit
In order to minimise the influence of measurement disturbances due to rotational accelerations and gravity gradients, the accelerometer sensor is positioned precisely in the centre of gravity of the CHAMP satellite (maximum deviation 2 mm). The exact orientation of the accelerometer axes are measured by means of star sensors which are attached to the thermal housing of the sensor unit. The thermal housing guarantees a thermal stability of better than 1°C per orbit for the accelerometer sensor in order to minimise thermal effects on bias and scaling factor.
The GPS Receiver TRSR-2 onboard CHAMP is provided by NASA and manufactured at NASA's Jet Propulsion Laboratories (JPL). In combination with the STAR accelerometer it serves as the main tool for high-precision orbit determination of the CHAMP satellite. Additional features are implemented for atmospheric limb sounding and the experimental use of specular reflections of GPS signals from ocean surfaces for GPS-altimetry. A synchronisation pulse delivered every second is used for precise onboard timing purposes, and the autonomously generated navigation information is used by both the CHAMP AOCS and the star sensors to update their orbital position.
The receiver has the following measurement modes:
CHAMP Black Jack GPS flight receiver
The low orbiting CHAMP satellite and each spacecraft of the high orbiting GPS satellite configuration establish a so called high-low satellite-to-satellite (SST) link. Each of the GPS S/Cs is transmitting a PRN modulated L1 and L2 signal which the TSRS-2 receiver onboard of CHAMP acquires for a maximum of 12 satellites at the same time. From these signals the orbiting receiver generates at a frequency of 0.1 Hz pseudo-ranges and carrier phases for all satellites which were in lock at this time instant. By using pseudo-ranges from at least 4 different GPS S/Cs with known ephemeris at the same time, both the three-dimensional coordinates of the CHAMP receiver and their respective change with time can be solved for, thus a navigation solution is obtained. The accuracy of this solution depends on the availability of the classified P-code inside the receiver and the satellite constellation used for coordinate determination and can range down to a few meters after post-processing. With the less precise but unclassified C/A-code an accuracy of several tens of meters is still obtainable after post-processing. By making use of the much more precise carrier phase tracking data and a simultaneous full dynamic solution for the orbits of the GPS satellites and the CHAMP spacecraft, corrections for reference frame and dynamic and measurement model parameters high accurcy results for CHAMP will be achievable. The remaining orbit perturbations are the input for the improvement of the gravity field modeling.
A new field of use for GPS measurements is the observation of radio occultation of GPS S/C by the Earth from a low-orbiting satellite. For this purpose the GPS signal of the satellite to be occulted is sampled with considerably higher frequency than usual (50 Hz instead of 0.1 ... 1 Hz) during crossing the lower layers of the Earth atmosphere in comparison to a non-occulted GPS S/C.
A downward positioned GPS antenna will be experimentally used to collect specular reflections of GPS signals from the ocean's surface. Knowing the precise position of the transmitting GPS S/C and of the CHAMP satellite as well, quasi-altimetry measurements can be carried out.
The TRSR is a sixteen-channel Global Positioning System receiver. Up
to twelve channels can be used for Precise Orbit Determination (POD). The
TRSR is completely self-initialising from a cold start. Once started, it
operates fully autonomously. The receiver decides which GPS S/C to track
in the various tracking modes, solves for the state vector (navigation
solution) for the CHAMP satellite, schedules occultation and altimetry
measurements and solves for the offset between the GPS receiver time and
the GPS time in order to provide an extremely accurate synchronisation
pulse for the CHAMP onboard subsystem.
| Computed position in telemetry | < 60 m |
| Time calibration accuracy | < 1µs
from GPS time
(resolution 0.1 ns) |
| Dual-frequency range and integrated carrier for POD at a 1s interval: | |
|
|
< 0.2 cm |
|
|
< 30 cm |
| Dual-frequency integrated carrier phase and amplitude for atmospheric occultation: | |
|
|
< 0.05 cm |
|
|
< 50 cm |
| Limb-sounding observables (prior to atmospheric de-focusing): | |
|
|
< 0.05 cm (1 s) |
|
|
< 0.15 cm (1 s) |
The GPS receiver system on CHAMP consists of a Receiver/Processor Assembly
(RPA) containing the RF down-converter sections, an internal bus and the
cold-redundant base-band processor cards, the RF coaxial cables to the
4 antennas and the antennas itself. The following antenna types will be
used for the TRSR GPS receiver on CHAMP: Zenith mounted POD antenna on
a choke ring Spare POD antenna on the aft pannel High-gain helix antenna
for occultation measurements, mounted on the aft pannel (20 inclined towards
nadir) High-gain nadir-viewing helix antenna for GPS altimetry experiments
The Laser Retro Reflector is a passive payload instrument consisting of 4 cube corner prisms intended to reflect short laser pulses back to the transmitting ground station. This enables to measure the direct two-way range between ground station and satellite with a single-shot accuracy of 1 ... 2 cm without any ambiguities. These data will be used for precise orbit determination in connection with GPS for gravity field recovery, calibration of the on-board microwave orbit determination system (GPS) and two-colour ranging experiments to verify existing atmospheric correction models. The Laser Retro Reflector was developed and manufactured inhouse at GFZ.
The Laser Retro Reflector (LRR) consists of only 4 cube corner prisms arranged in a densely packed array in order to have for most of the time only one prism contributing to the reflected signal. Even in the short observation periods where more than one prism contributes to the signal (e.g. near culmination of the CHAMP satellite) the structure of the reflected pulses caused by the LRR (signature) will not be resolved by the presently existing Satellite Laser Ranging (SLR) systems. The resolution of the range measurements to the LRR will depend exclusively on the SLR ground station hardware thus making the CHAMP LRR an ideal target for high-resolution two-colour ranging experiments.
The individual cube corner prism of the reflector are made from fused quartz with metal coated reflecting surfaces. Two of the dihedral angles are rectangular to high precision whereas one angle has an offset in order to produce a two-spot far-field pattern to deal with the aberration correction. The front faces of the prisms are uncoated and slightly spherical.
Optical design parameters of the CHAMP LRR:
| Vertex length | 28.0 ± 0.2 mm |
| Free aperture of the front face | 38.0 mm |
| Dihedral angle offset | -3.8 arc sec (less than 90°) |
| Radius of curvature of the front face | + 500 m (convex) |
| Index of refraction @ 532 nm | 1.461 |
| Distance of the two peaks of the far field | 24 arc sec |
| Width of the far field peaks (FWHM) | 6 arc sec |
| Width of the far field peaks at 20% of maximum | 10 arc sec |
The 4 cube corner prisms of the reflector array are mounted on a regular 45° pyramid.
The LRR is mounted on a bracket exactly below the satellite's centre of mass. In the nominal orientation the vector from the centre of mass to the reference point is directed to nadir (the CHAMP +Z-axis). Its length is 250 mm. This value is obtained as the sum of the mechanical structure carrying the reflector and the distance of the reference point from the LRR mounting surface. The distance between the CHAMP CoG and the LRR mounting surface is 256 mm.
As the reference point of the LRR array, the crossing point of the optical axes of all cube corner prisms is defined. This reference point is located outside of the structure of the reflector with a distance of 6.2 mm from the mounting plane. Thus, the total distance of the LRR reference point to the satellite CoG in +Z-direction is 250 mm.
For the computation of the exact value of the measured laser ranges
to the CHAMP LRR, one has further to add the scalar product of the vector
from the CHAMP CoG to the LRR reference point (250 mm in nominal orientation)
with the unit vector of the light direction. For this purpose the attitude
of the satellite has to be known. Because of the active attitude control
of the CHAMP satellite with a dead band of several degrees in all 3 axes,
the attitude data as delivered by the on-board star sensors shall be introduced
for highly precise data processing.
The Fluxgate Magnetometer (FGM) was developed and manufactured under contract by the DTU (Technical University of Denmark) Lyngby. The design is based on the CSC (Compact Spherical Coil) sensor which was newly developed for the Ørsted mission and presently demonstrates its outstanding performance in orbit. The FGM is probing the vector components of the Earth magnetic field. It therefore is regarded as the prime instrument for the magnetic field investigations of the CHAMP mission. The interpretation of the vector readings requires the knowledge of the sensor attitude at the time of measurement. For that reason the FGM is mounted rigidly together with star cameras (cf. Advanced Stellar Compass) on an optical bench. For redundancy reasons a second FGM is accommodated on the optical bench, 60 cm inward from the primary sensor.
The operational principle of fluxgate magnetometers is well know and has proven in many missions to be very reliable and to offer high performance. Special in this mission is the design of the sensor. Three orthorgonal sets of coils are wound on the surface of a 82 mm diameter sphere in a configuration which generates a homogeneous field within the whole spherical volume. The current through these coils are controlled by a feedback loop trying to cancel the ambient magnetic field in the interior. Three ring core sensors in the centre act as null-indicators. The particular design of the CSC has proven to exhibit superb linearity figures, no sensitivity to transverse fields and an excellent angular stability.
The FGMs cover the full range of the Earth's field, ±65 000 nT,
in all three components. The analogue outputs are digitised by 24 bit ADCs
achieving a quantisation step size of 10 pT. Deviations from linearity
are found to be in the range of ±100 pT and the overall noise level
is of the order of 50 pT (rms). In the nominal operation mode the field
vector is sampled at a rate of 50 Hz providing a spatial resolution along
the orbit of approximately 150 m. There are other modes which allow to
reduce the demands on data transmission. Both options, data compression
and reduced sampling rates can be freely combined.
| Range | ±65 000 nT |
| Resolution | 10 pT |
| Deviation from linearity | ±100 pT |
| Noise level | < 100 pT (rms) |
| Sample rate | 50 Hz (nominal), 10 Hz, 1 Hz |
| -3 dB bandwidth | 13 Hz |
| Offset drift | < 0.5 nT |
| Sensor
weight dimensions |
350 g (each) Ø 82 mm |
| Electronics
box (for both sensors) weight |
3.5 kg |
| Power Consumption
dimensions |
2 W (each)
204x194x101 mm3 |
Both CSC sensors are mounted together with the star cameras (ASC) on a common optical bench providing a mechanical stability between these systems of better than 10 arcsec. The optical bench as a part of the boom is placed about halfway between the satellite body and the Overhauser Magnetometer (OVM) at the tip. This location is a compromise between avoidance of magnetic interference from the spacecraft and cross-talk between FGM and OVM.
Flying two FGMs on CHAMP is mainly
for redundancy reasons, but such an arrangement with the sensors separated
by 60 cm can also be used as a gradiometer when both instruments are operated.
Differences in the reading of the two instruments can effectively be employed
to detect and localise magnetic disturbances from the spacecraft.
The Overhauser Magnetometer (OVM) was developed and manufactured under contract by LETI (Laboratoire d'Electronique de Technologie et d'Instrumentation) at Grenoble. It serves as the magnetic field standard for the CHAMP mission. The purpose of this scalar magnetometer is to provide an absolute in-flight calibration capability for the FGM vector magnetic field measurements. A dedicated program ensuring the magnetic cleanliness of the spacecraft allows for an absolute accuracy of the readings of < 0.5 nT.
The underlying idea of this scalar magnetometer is based on the principle of proton magnetic resonance. If a proton-rich liquid is exposed to a DC magnetic field, the protons will start to precess around the field direction with a frequency strictly proportional to the applied field magnitude. In principle there is no dependency on field direction, on temperature and no drift. By exactly measuring the precession frequency (0.8 - 3 kHz) an absolute figure of the ambient magnetic field strength can be derived. The employed proportionality constant between frequency and field strength, called gyro magnetic ratio, will be one of the important CHAMP Standards.
The OVM samples continuously the ambient field strength at a rate of
1 Hz. It can cope with fields from any direction. There are no dead zones
as often encountered with instruments of this type. The deviation from
omni-directionality is less than 0.2 nT. The internal crystal oscillator
is regularly checked against GPS clock to ensure a precise determination
of the proton precession frequency.
| Range | 18 000 - 65 000 nT |
| Resolution | 10 pT |
| Noise level | < 50 pT (rms) |
| -3 dB bandwidth | 0.28 Hz |
| Sample rate | 1 Hz |
| Absolute accuracy | < 0.5 nT |
| Sensor weight | 1 kg |
| Electronics weight | 2 kg |
| Power consumption | 4.5 W |
| Sensor dimensions | Ø 90x180 mm |
| Electronics box dimensions | 200x135x76 mm3 |
In order to keep the influence of the DC magnetic stray field of the
spacecraft as low as possible the OVM sensor is mounted at the tip of a
4 m long deployable boom. The electronics box is placed inside the satellite
body to provide more comfortable environmental conditions.
The Advanced Stellar Compass (ASC) has been developed and fabricated under contract by the DTU (Technical University of Denmark, Lyngby). The design of this star imager is based on a new development presently flown on the Ørsted satellite. On CHAMP there are two ASC systems each consisting of two Camera Head Units (CHU) and a common Data Processing Unit (DPU). One ASC is part of the magnetometry optical bench unit on the boom and the other provides high precision attitude information for the instruments fixed to the spacecraft body. Additionally the ASCs serve as sensors for the satellite attitude control system.
A special feature of the ASC, very important for magnetometry missions, is the magnetic cleanliness of the camera heads. It allows to mount the camera and the magnetometer close together on a rigid structure which significantly improves the validity of attitude solutions transferred from one system to the other. The ASC is a fully autonomous system that directly outputs the attitude quaternions in a preselectable coordinate system.
An image of the stars within the field-of-view is acquired by integrating the light focused onto a photo-sensitive charge coupled device (CCD) array. The pattern is serially read out, digitised and fed to a microprocessor. The digital image is than sifted for stars brighter than mV = 6 and corrected for lens distortions resulting in a list of calculated star centroids with sub-pixel precision. The determined star centroids are subsequently matched against real star positions derived from an on-board HIPPARCOS star catalogue. The star attitude which fits best the observed pattern is output result. For further improvement of the attitude solution a GPS updated orbit model is maintained to correct for astronomical aberration and also the epoch of star constellation is taken into account.
The ASC on the boom provides the high attitude precision needed for the magnetic field vector measurements. The two transverse directions, Right-Ascension and Declination, can be determined with an accuracy in the arc second range. Whereas the rotation angle about the boreside is poorer by about a factor of 5. An improvement of this situation can be achieved by combining the readings from the two sensor heads which are separated by an inter-boresight angle of about 90°, in a single solution and disregarding the poorer Rotation readings.
The ASC on the spacecraft body provides attitude data primarily for the three component STAR accelerometer and the Digital Ion Drift Meter. This information is however also required for the proper reduction of the GPS data, Laser ranging data and the attitude control.
Due to the nadir orientation of CHAMP the Sun will blind one or two
of the CHUs over certain parts of the orbit. The camera heads can tolerate
the full Sun in the optic for an unlimited time. The attached baffles provide
for an Sun exclusion angle of about 30°. Having the Moon in the field-of-view
does not influence the attitude determination.
| Attitude determination precision | 4 arcsec (3s , BOL) |
| Field of view | 18.4 ° x 13.4° |
| Sampling rates | 1 Hz (nominal), 0.5, 2 Hz |
| Magn. moment CHU | 10-5 A/m2 |
| Power consumption | 8 W |
| CHU weight | 200 g (exclusive baffles) |
| DPU weight | 800 g |
| CHU dimensions | 50x50x45 mm3 |
| DPU box dimensions | 100x100x100 mm3 |
The body-mounted CHUs are rigidly fixed to the thermal housing of the accelerometer comprising a 90° inter-boresight angle (IBA) which is determined by the satellite design with its side solar panels being tilted 45° versus the horizontal instrument platform. For the boom optical bench an 102° IBA has been chosen in order to provide optimum simultaneous usage of both camera heads (i.e. minimise blinding of both sensors depending on the satellite orbit constellations wrt. Sun, Earth and Moon).
Magnetic cleanliness, size and weight of the CHU make this ASC ideal for use in conjunction with magnetometry instruments on a boom.
All CHUs are equipped with an inner baffle
(12cm) and an outer 25cm long baffle. This double-stage baffle is chosen
to optimise the Sun exclusion angle for the given baffle length.
The Digital Ion Drift Meter (DIDM) is provided by the AFRL (Air Force Research Laboratory, Hanscom). The DIDM is an improved version of an analogue ion drift-meter type flown successfully on many upper atmospheric satellites. The purpose of this instrument is to make in-situ measurements of the ion distribution and its moments within the ionosphere. A number of key parameters can be determined from the readings, such as the ion density and temperature, the drift velocity and the electric field by applying the (v x B)-relation. Together with the magnetic field measurements these quantities can be used to estimate the ionospheric current distribution. Knowing these currents will help significantly to separate internal from external magnetic field contributions.
In combination with DIDM a Planar Langmuir Probe (PLP) is operated. This device provides auxiliary data needed to interpret the ion drift measurements. Quantities that can be derived from the PLP sweeps are spacecraft potential, electron temperature and density.
The sensing elements on DIDM are two ion detectors mounted side-by-side onto the Data Processing Unit (DPU). Ions entering the pin-hole aperture fall into a Retarding Potential Analyser (RPA) cup. Those ions with energies surmounting the settable potential barrier will be guided by an electric field onto a Multi-Channel Plate (MCP). Using an anode which is divided into a 16x128 pixel array the impact location can be detected. By taking into account the ion optics the incoming direction of the particle can be tracked back. Accumulating many ion impacts provides an image of the ion distribution in the velocity space. The cross-track components of the ion velocity are well determined by the location of the spot on the anode. The along track component has either to be deduced from the effect of varying RPA voltages or from a comparison of the results obtained by the two detector heads. They are both skewed by 10° with respect to the nominal flight direction giving a spread angle of 20°.
The ion velocities detected by DIDM are dominated by the satellite orbital velocity. To obtain rest frame results great care has to be taken to consider the spacecraft velocity vector correctly. It is foreseen to derive precise orbit data from the GPS readings and high quality attitude data from the ASC.
Ion temperatures can be estimated from the width of the distribution obtained on the anode. The total count rate is a measure for the ion density. For checking the actual distribution images of the pixel map can be returned. One can selected between partial images centred about the peak count or a full image.
The PLP voltage is swept for 1 sec every 15 sec typically between -2.5 and +2.5 V in 32 steps. A selectable bias voltage can be added to account for the S/C potential. By interpreting the measured current/voltage characteristic the above mentioned plasma parameters can be determined. The spacecraft floating potential is measured during the remaining 14 sec.
The performance of the DIDM instrument has been verified by tests in a plasma chamber. Each detector resolves a 35° half-cone of space. The pixel sizes allow for a gross resolution of about ±1° which can be improved by fitting the results to a theoretical distribution function. Special look-up tables will be used to correct for ion optic distortions.
The instrument provides a large variety of selectable modes. Drift measurements can be obtained at rates from 1 to 16 Hz. There are 8 selectable sets of RPA voltage sweeps. They are designed to allow for optimising the measurement modes, e.g. for fast sampling, high resolution, mass resolution etc. Both detectors can be operated fully independently. This provides some kind of redundancy but generally will be used to enhance the scientific return.
The prime purpose of the PLP is to determination
the spacecraft potential. An error of 0.1 V in this quantity is equivalent
to an along-track velocity of 75 m/s. The PLP is designed to measure currents
drawn from the ambient plasma down to 5 nA.
| Range of ion density | 108 - 1012 ions/m3 |
| Range of ion temperature | 200 - 55 000 K |
| Range of drift velocity | 0 - 6 km/s |
| Range of electric field | 0 - 300 mV/m |
| Resolution of ion velocity | < 1° direction, < 130 m/s speed |
| Resolution of electric field | < 4 mV/m |
| Sample rates
DM mode RPA mode PLP mode |
0, 1, 2, 4, 8, 16 HZ 0, 8, 16 Hz 0, 1/15 Hz |
| Power consumption | 5 W |
| Weight | 2.2 kg |
| Dimensions | 153x150x109 mm3 |
The DIDM instrument is mounted on a bracket
outside the spacecraft body. The two detectors have to point into the ram
direction for proper operation. An accommodation as far down as possible
is chosen to minimise the wake effect caused by the boom. The slanted face
sheet behind DIDM is well conducting (Aluminium). By pointing into the
ram direction it shall provide a close electrical contact of the spacecraft
to the ambient plasma avoiding a charge-up of the satellite.