Minimize LISA Pathfinder

LPF (LISA Pathfinder) Mission

The LISA Pathfinder mission of ESA (formerly the mission was called SMART-2) is a technology demonstration mission for LISA, a kind of physics research laboratory in space, with the objective to test and verify the key technologies needed for highly accurate formation flying and precise measurement of the separation (metrology) between two very distant spacecraft.1) 2) 3) 4) 5) 6)

The LISA Pathfinder mission will test in flight:

- Inertial sensors

- Interferometry between free floating test masses

- Drag Free and Attitude Control System (DFACS)

- Micro-Newton propulsion technology: FEEP (Field Emission Electric Propulsion) and colloidal thrusters of NASA/JPL.

The basic idea of LISA Pathfinder is to squeeze one arm of the LISA constellation from 5 million km to a few tens of cm!


Figure 1: LISA Pathfinder experiment concept (image credit: ESA)

Legend to Figure 1: The top left image shows the classical Einstein thought experiment to measure the spacetime curvature. This is the basis for all gravitational wave detectors, e.g. LISA (top right). LISA Pathfinder will not only pave the way for LISA, but will also demonstrate the main assumption of the thought experiment: that free particles follow geodesics.


Some background on the LISA (Laser Interferometer Space Antenna) mission:

LISA Pathfinder is a necessary precursor mission for LISA, the joint ESA and NASA formation-flying mission (launch scheduled for 2018) of three spacecraft. The goal of the LISA fundamental physics mission is to detect low-frequency gravitational waves in the range of 10-4 to 10-1 Hz (according to Einstein's Theory of General Relativity the force of large gravity changes, generated by a collapsing star or an entire galaxy, will produce tiny ripples of gravity waves) - requiring technologies that have so far never been tested. With the detection of these waves, which for example originate from black holes in close orbit around each other, an entirely new window to the universe will be opened.

In this concept, the three LISA spacecraft constellation form an equilateral triangle (a giant interferometer) with an armlength of 5 million km, inclined by 60º against the ecliptic (Figure 2). Plans call for LISA's trio of spacecraft to orbit the sun at the same distance as Earth, but trailing about 50 million km in orbit behind our planet (representing an angle of about 20º with respect to the Earth-Sun direction). 7) 8)


Figure 2: Orbital configuration of the LISA mission concept (image credit: ESA)

The concept of detecting a gravitational wave with an interferometric configuration like LISA is being realized using a transverse and traceless (TT) coordinate system. The goal is to detect the “ripples of gravitational waves” by the measurement of the time variation that can be detected by the giant interferometer.

In the LISA measurement concept, each spacecraft houses two proof masses; changes in the distance between the proof masses are measured interferometrically to a level of 10 pm. To insure that the proof masses can follow as close as possible a purely gravitational orbit, their position inside the spacecraft is constantly being monitored. Capacitive and interferometric sensors determine the position of the spacecraft with respect to the proof masses. A drag-free system and proper shielding must counteract the disturbance forces of the orbiting spacecraft. 9) 10) 11)


Figure 3: Schematic illustration of the LISA interferometer concept (image credit: ESA, NASA)

Due to the required level of residual forces on the proof masses, LISA faces a number of technological challenges, such as tight requirements on magnetic cleanliness, thermal stabilization, charge management, and above all the internal gravitational balancing to minimize the gravity gradient forces on the proof masses. The most critical technologies, such as the drag-free system (or the disturbance reduction system) and the interferometry are planned to be addressed by the precursor mission LISA Pathfinder - with the ESA payload LTP and the NASA payload DRS.


Figure 4: Artist's view of LISA's yearly orbit around the Sun. The rotation of the triangle that the spacecraft form can be seen in the picture (image credit: ESA)


Relationship between LISA Pathfinder and LISA

From the outset, the LISA Pathfinder mission has been designed such that the technology can be directly transferred to LISA with little, or no, changes. However, as LPF is a pathfinder, and in order to keep costs down, the Level One mission performance requirement is relaxed with respect to the LISA hardware requirements. Specifically, the requirement on the relative acceleration noise of the test masses is relaxed by an order of magnitude in performance, and by a factor of thirty in frequency (Figure 5). 12) 13)


Figure 5: Relative acceleration noise requirements of the LISA Pathfinder and LISA test masses (image credit: NASA, ESA)

Legend to Figure 5:

• The graph on the left side is showing the Level One acceleration noise requirement of LISA Pathfinder and LISA. The line labelled LISA Pathfinder Requirement shows the value as listed in the Science Requirements Document. The line labelled LPF CBE shows the Current Best Estimate of the expected performance of the mission. The gap between the LPF CBE line and the LISA Requirement represents the extrapolation required to transfer the LPF technology to LISA.

• The graph on the right side is showing the Level One Displacement Noise requirements. As can be seen, the LISA Pathfinder requirement on the performance of the readout interferometer is approximately equal to the performance of the LISA local (test mass readout) interferometer. The line labelled LPF CBE is the measured performance of the EM interferometer (optical bench, laser, phase meter).

The relaxation in the acceleration noise requirement significantly reduces the environmental requirements levied on the spacecraft. In particular, this relaxation is most evident on the thermal stability, gravitational balancing and magnetic cleanliness of the spacecraft. The environmental relaxation allows the LPF spacecraft to be in orbit around L1, as opposed to LISA's heliocentric Earth-trailing orbit. This has two advantages, namely in the time required to reach the desired orbit, and more importantly, it greatly reduces the communication requirements onboard the spacecraft (distance to LPF is approximately 1:5 million km as opposed to 50 million km for LISA).


LISA Pathfinder



LPF status


35 mW @ 1064 nm

2 W @ 1064 nm

LISA will use LPF-like laser as its master oscillator. Higher power is achieved by optical fiber amplifier

FM in final testing

Laser frequency stabilization

Unequal arm Mach-Zehnder

Unequal arm Mach-Zehnder or reference cavity

The laser onboard LISA requires several stages of frequency correction. The first stage, pre-stabilization, could adopt an LPF type unequal arm-length Mach-Zehnder interferometer as the frequency discriminator

Fully tested using EMs of optical bench, laser and phasemeter


AOM (Acousto-Optic Modulator)

EOM (Electro-Optic Modulator)

No demonstration of LISA electro-optic modulator on LPF

FM (Flight Model) in final testing

Optical bench

Hydroxy catalysis bonded Zerodur bench

Hydroxy catalysis bonded Zerodur bench

Demonstration of the LISA optical bench manufacturability and pathlength was one of the main technology developments in LPF. Several technologies have been developed including hydroxy-catalysis bonded ultra-stable optical bench, and quasi-monolithic fiber injector assemblies

FM under construction


SBDFT (Single Bin Discrete Fourier Transform) algorithm

Digital PLL (Phase-Locked Loop)

LISA requires a high-frequency phasemeter due to the large Doppler shifts between the S/C. As LPF uses a significantly lower heterodyne frequency, it does not require such a sophisticated phasemeter

FM under construction

Inertial Sensor (IS)



The inertial sensor is the main component of LISA which cannot be tested on the ground. The demonstration of the inertial sensor performance is the main reason for flying LPF. The LPF inertial sensor has been designed as the LISA inertial sensor from the beginning


IS-Test Mass

46 mm Au:Pt cube

46 mm Au:Pt cube

The LPF test mass is identical to the LISA test mass

FMs delivered

IS-Electrode Housing

Molybdenum housing with gold-coated sapphire electrodes

Molybdenum housing with gold-coated sapphire electrodes

LISA electrode housing will be identical to the LPF electrode housing

FM replica delivered for testing. Flight units under construction

IS-Caging Mechanism

3 actuator design

3 actuator design

LISA caging mechanism will be identical to the LPF caging mechanism. It consists of the launch lock CMSS (Caging Mechanism Support Structure) and GPRM (Grabbing, Positioning and Release Mechanism).

GPRM FMs delivered. CMSS FMs under construction

IS-Vacuum enclosure

Titanium enclosure with getter pump

Titanium enclosure with getter pump and gate valve

The basic vacuum enclosure of LISA will be identical to the LPF vacuum enclosure with the exception of a gate valve which can be used to vent the interior to space

FMs under construction

Front End Electronics (FEE)

Differential Capacitive Bridge

Differential Capacitive Bridge

Due to frequency relaxation, LPF FEE performance has not been demonstrated at LISA's lowest frequency band. In some cases, LPF FEE is more difficult than LISA due to in-band actuation along the sensitive axes.

FMs in final testing

Charge management

Photoelectric discharge

Photoelectric discharge

Only change could be utilization of solid state UV light source as opposed to gas discharge lamp.

FMs delivered

Micro Newton Thrusters

FEEPs (Field Emission Electric Propulsion)/Colloids


The demonstration of µN thrusters is one of the primary goals of the LPF mission. Both FEEP and colloid thrusters will be demonstrated. The results of LPF will determine which thruster will be chosen for LISA. LPF FEEP thrusters have been designed to meet the LISA lifetime requirements, although full LISA lifetime will not be demonstrated with LPF

Cs FEEP FMs under construction




Additional DoF (Degree of Freedom) comes from constellation breathing (not applicable to LPF).

Open loop test complete.
Closed loop test ongoing



Off-axis Schiefspiegler

LPF does not carry a telescope


Table 1: Relationship between the LISA technology and LISA Pathfinder, LPF status as of April 2009 (Ref. 12)



LISA Pathfinder Mission

LISA Pathfinder has been introduced to mitigate the risks of the LISA mission. The main goal of the LISA Pathfinder mission is to demonstrate the concept of the gravitational wave detection using a single spacecraft: it will put two test masses in a near-perfect gravitational free-fall and control and measure their motion with unprecedented accuracy. This is achieved through state-of-the-art technology comprising the inertial sensor system, the laser metrology system, the drag-free control system and an ultra-precise micro-propulsion system.

Major mission objectives are: 14) 15) 16)

• LISA Pathfinder's experiment concept is to prove geodesic motion by tracking two test masses nominally in free fall through laser interferometry with picometer (10-12 m) distance resolution. LISA Pathfinder will show that the relative parasitic acceleration between the masses, at frequencies around 1 mHz, is at least two orders of magnitude smaller than the value demonstrated so far or to be demonstrated by any planned mission.

• To demonstrate drag-free and attitude control in a single spacecraft with two proof masses

• Test the feasibility of laser interferometry at the level of accuracy envisaged for LISA

• Test endurance of the different instruments and hardware in the space environment.


The basic elements to achieve and prove geodesic motion are the following:

- Free floating test masses equipped with motion sensors in all degrees of freedom and free of dynamical disturbances (< 3 x 10-14 m s-2 Hz1/2 @ 1 mHz)

- Low-thrust (~10 µN), low-noise (0.1 µN / Hz ) proportional thrusters to push the spacecraft to follow the test masses

- A high resolution laser interferometer to measure test mass relative displacement, 18-degree of freedom dynamical control laws

- Gravitationally “flat” (< 5 x 10-11 g) and gravitationally stable spacecraft to host the test masses.


Test masses

Test mass environment

Tracking method


Test-masses geodesic motion performance
(m s-2 Hz-1/2 @ 1 mHz)

Drag-free (Residual S/C acceleration)
(m s-2 Hz-1/2 @ 1 mHz)


Accelerometer test-masses (<100 g)

< 200 µm gaps from
electrodes. Mechanical contact via grounding wire

Radio link plus capacitive sensing





Differential accelerometer test masses (< 0.5 kg)

~ 200 µm gaps from electrodes. Mechanical contact via grounding wire

Capacitive sensing relative to S/C


2 x 10-12

3 x 10-10


Accelerometer testmasses (320g)

~ 300 µm gaps from electrodes.
Mechanical contact via grounding wire

Capacitive sensing relative to S/C


3 x 10-12

3 x 10-8


Gravity Reference Sensor test masses (Au-Pt, 2 kg)

No mechanical contact. 4 mm gaps

High resolution TM-TM interferometry

Interplanetary (L1), drag-free

3 x 10-14

3 x 10-13

Table 2: Comparison of main features of missions requiring geodesic motion


Some background on the history of the various LPF mission development stages: 17)

• The LPF (LISA Pathfinder) mission was initially proposed in 1998 under the name ELITE (European LISA Technology Experiment). A homodyne interferometer was to be flown. The planned launch date of ELITE was 2002.

• The ELITE proposal was further refined over the next two years, and in 2000 was submitted to ESA as SMART-2 (Small Missions for Advanced Research in Technology-2). The initial design of the SMART-2 mission comprised two formation-flying spacecraft to serve a LISA Pathfinder and also a Darwin Pathfinder payload objective.
However, after further studies, it was decided to focus the mission on a single spacecraft demonstrating drag-free control dedicated to the future LISA mission. The planned launch date of SMART-2 was 2006. 18)

• The renaming of the SMART-2 mission to “LISA Pathfinder” occurred in 2004 to account for the changed mission objectives (technology demonstrations for the LISA mission only). In October 2004, the ESA Science Program Council (SPC) approved the LTP Multi-Lateral Agreement, detailing the national agency responsibilities for the construction of the LTP.

• Originally, LISA Pathfinder/SMART-2 consisted of two payloads, namely LTP (LISA Technology Package) provided by ESA member states, and the NASA-provided DRS (Disturbance Reduction System), also known as ST7 (Space Technology 7). Each payload consisted of two gravitational reference sensors (GRS), an interferometric readout system, drag-free and attitude control system (DFACS), and micro-Newton thrusters.
In October 2005, the NASA-provided DRS was descoped; the DRS now consists of micro-Newton thrusters and DFACS, and will rely on the LTP as its gravitational reference sensor.

• In February 2006, after the successful completion of the Mission Preliminary Design Review, LISA Pathfinder entered the Development Phase.

• LTP (LISA Technology Package) successfully passed CDR (Critical Design Review) in November 2007

• LPF (LISA Pathfinder) Mission CDR in December 2008.

• LPF STOC ()Science and Technology Operations Center) passed CDR in Sept 2009.

• A launch of LPF is scheduled for 2014 on a Vega launcher of ESA.




The LISA Pathfinder spacecraft is being built and integrated by EADS Astrium Ltd. of Stevenage, UK (contract award in June 2004). The spacecraft is comprised of the science spacecraft (science craft) and a separable propulsion module for apogee raising. The science spacecraft contains the two main instruments LTP and DRS, and is covered with one single fixed solar array. The mass of the science craft after arrival at the operational orbit is about 480 kg. The sciencecraft dimensions are: 2.1 m diameter and 1.0 m in cylinder height. The propulsion module has a mass of 1420 kg (including fuel). The total launch mass is about 1900 kg (S/C dimensions of 2.9 m in length and 2.1 m diameter). 19)


Figure 6: Artist's rendition of the LISA Pathfinder spacecraft in orbit (image credit: ESA)

The extremely stable and lightweight structure of the science module of LISA Pathfinder is made of CFRP (Carbon Fiber Reinforced Plastic) sandwich panels and shells. The mating ring of the propulsion module is made of aluminum alloy. The structure will provide a stable mounting to LISA Pathfinder's gravitational sensor technology package (LTP and DRS), limiting deformations at the interface during flight to less than 1 x 10-8 m Hz-1/2 in the instrument's measurement bandwidth between 1 and 30 mHz.


Figure 7: Photo of the the LISA Pathfinder science module structure (image credit: ESA)


Figure 8: LISA Pathfinder Flight Model in launch configuration installed on the sine vibration shaker of IABG in Ottobrunn, Germany (image credit: ESA)

Legend to Figure 8: The sine test is performed (spring 2011) on the spacecraft in launch configuration with the PRM (Propulsion Module) mated to the SCM (Science Module) on top. 20)


Launch: A launch of the LISA Pathfinder spacecraft is scheduled on a Vega vehicle (as one of the VERTA flights) for 2015 from Kourou.

Orbit: A halo orbit about the Lagrangian point L1 will be the operational orbit (1.5 million km away from Earth in the direction toward the sun). This location helps to minimize disturbances from the Earth's gravity, magnetic field and the atmosphere.

The spacecraft and the propulsion module will be injected into a slightly elliptical parking orbit of about 200 km x 1600 km at an inclination of 63º. From there, it will use the propulsion module (with a series of apogee burns) to enter into a transfer orbit and reach eventually its operational orbit around L1. After the transfer orbit the propulsion module separates from the science spacecraft (Figure 9). The LISA Pathfinder spacecraft will be stabilized using the micro-Newton thrusters, entering a Lissajous orbit around L1 (500,000 km x 800,000 km orbit around L1). 21) 22)

Following the initial on-orbit check-out and instrument calibration, the in-flight demonstration of the LISA technology will then take place. The nominal lifetime of the mission is 180 days; this includes the LTP operations (90 days), the DRS operations (60 days), and a period of joint operations (30 days) when the LTP will control the DRS thrusters.

The constant orientation in the Earth-Sun direction of the spacecraft will provide a stable thermal environment.


Figure 9: Concept of the LISA Pathfinder orbit injection sequence (image credit: ESA)

RF communications: X-band links will be used for TT&C as well as for science data service functions. The ground interface consists of a single 35 m X-band antenna providing about 8 hours of communications per day. Spacecraft operations will be done at ESOC while science and technology service functions will be provided by ESAC.

Communications to the spacecraft will be performed in X-band through a network of ground stations, including Kourou, Maspalomas and Perth, during LEOP (Launch and Early Operations Phase).

A new class of X-band transponder, referred to as X2PND, has been designed and developed by TAS-I (Thales Alenia Space – Italia) for LISA Pathfinder and for the Gaia mission of ESA. The compact X2PND device has a mass of 3.1 kg (including the diplexer), achieved through a high degree of integration. 23)


Figure 10: Top-level block diagram of the X2PND device (image credit: TAS-I)


Figure 11: Photo of the LISA Pathfinder launch composite at the IABG test facility in Munich, Germany (image credit: ESA) 24)



Sensor/experiment complement: (LTP, DRS)

The sensor complement comprises two packages: the LISA Technology Package (LTP) and the Disturbance Reduction System (DRS). Both will test the key technology of drag-free control by means of proof masses. 25)

Metrology concept: In LISA Pathfinder, the traditional distinction between spacecraft and payload disappears. The instrument really involves the entire spacecraft. It is fair to state that LISA Pathfinder implements a “formation flight” of three orbiting bodies, namely the spacecraft and the pair of test-masses. This formation flight is implemented using one of several variants of the basic “drag-free” control scheme: One of the test-masses is in pure free-fall in all translational degrees of freedom (x, y, and z in Figure 12) and no force is purposely applied to it. The spacecraft follows this first test-mass within a standard drag-free control scheme in all translational degrees of freedom. The second test-mass is free along y and z and the spacecraft can follow this by using rotation around z and y, respectively. 26) 27) 28) 29)


Figure 12: Metrology concept of the LISA Pathfinder mission (image credit: ESA)

The LISA Pathfinder measurement scheme, with the separate high-resolution optical readout of test-mass motion relative to the spacecraft, allows test-mass to test-mass tracking with accuracy unspoiled by the spacecraft motion even for test-masses located in different spacecraft. Indeed, as in LISA, one can track one test-mass relative to its hosting spacecraft and then one spacecraft relative to the other one and then reconstruct the test-masses' relative motion by adding up these three measurements. Thus LISA Pathfinder demonstrates the possibility of undertaking high resolution geodesy with test-mass to test-mass tracking.


LTP (LISA Technology Package)

LTP is a payload developed for ESA by the European scientific community using national funding (Table 3). Institutes and industries of the following countries are involved: France, Germany, Italy, Spain, Switzerland, The Netherlands, and the United Kingdom. EADS Astrium GmbH (Friedrichshafen, Germany) is the prime contractor for LTP payload integration (subsystem deliveries from national teams). LTP co-Principal Investigators are Stefano Vitale of Trento University, Italy and Karsten Danzmann of the Max-Planck-Institut für Gravitationsphysik (Albert Einstein Institut) in Hannover, Germany.




Space Agency


APC (Laboratoire Astroparticule et Cosmologie), University of Paris,
Oerlikon Space (Switzerland)

Laser Modulator



MPI-AEI (Albert Einstein Institut), Hannover,
Astrium GmbH, Friedrichshafen
Tesat Spacecom, Backnang
ZARM, Bremen
Kayser Threde

Karsten Danzmann, PI: Interferometer design
LTP Architect
Reference Laser Unit (RLU)
Drag-Free and Attitude Control
Laser Assembly



University of Trento,
CGS (Carlo Gavazzi Space)
Thales Alenia Space

Stefano Vitale, PI: Inertial Sensor Design
Inertial Sensor Subsystem
Test Mass, Electrode Housing
Micro-Newton FEEP Thruster


The Netherlands

SRON (Space Research Organization Netherlands)

ISS-SCOE (Special Check Out Equipment)



IEEC (Institut d’Estudis Espacials de Catalunya) / University of Barcelona

DMU (Data Management Unit), Data Diagnostic System

Spanish National Space Program


ETH Zürich/Oerlikon Space

ISS Front End Electronics (FEE)

Swiss Space Office

United Kingdom

University of Birmingham
University of Glasgow
Imperial College London, RAL/STFC
Astrium Ltd.

Phasemeter Assembly
Optical Bench Interferometer
Charge Management System
Optical Bench prime contractor



Thales Alenia Space, Italy
Astrium GmbH, Germany

Caging Mechanism
LTP Architect and System Engineer for LTP


Table 3: Responsibilities in the manufacture of the LISA Technology Package

The LTP concept on LISA Pathfinder represents one arm of the LISA constellation interferometer, in which the distance between the two proof masses is reduced from 5 million km to about 35 cm. As in LISA, the proof masses fulfil a double role: they serve as mirrors for the interferometer and as inertial references for the drag-free control system.

The mission goals for the LTP are: 30) 31) 32) 33) 34) 35)

• To demonstrate drag-free and attitude control in a spacecraft with two proof masses in order to isolate the masses from inertial disturbances.

- The LISA Pathfinder LTP required performance is: ≤ 3 x 10-14 ms-2 Hz-1/2 in the bandwidth 10-3 to 10-1 Hz along the sensitive axis. This is a factor of ~ 7 larger than what is required in LISA.

- The LISA LTP required performance is: ≤ 10-15 ms-2 Hz-1/2 in the bandwidth 10-3 to 10-1 Hz along the sensitive axis.

• To demonstrate the feasibility of performing laser interferometry in the required low-frequency regime with a performance as close as possible to 10-12 ms-2 Hz-1/2 in the frequency band 10-3 to 10-1 Hz , as required for the LISA mission

• To assess the longevity and reliability of the capacitive sensors, thrusters, lasers and optics in the space environment.

• The final objective of LISA Pathfinder is to confirm the overall physical model of forces that act on a test-mass in interplanetary space.

The LTP measurement scheme is shown in Figures 13 and 14. LTP contains the two (partially) free-floating TM (Test Masses), each surrounded by a sensor cage. Each cage is rigidly attached to the optical bench, i.e. the spacecraft. The distance between the two test masses is measured with an optical metrology system (laser interferometer) along the sensitive x-axis which is the nominal line of connection between the two test masses. The proof (or test) masses are made of a gold-platinum, low magnetic susceptibility alloy, have a mass of m = 1.96 kg and are separated by a nominal distance of 376 mm. The proof masses for LISA Pathfinder are the same as those foreseen for LISA.


Figure 13: Test mass degrees of freedom (image credit: EADS Astrium GmbH)

Metrology system (Figure 14): The optical bench of LTP in LISA Pathfinder uses a total of four interferometers:

- 1) one to measure the distance between the proof masses,

- 2) one to measure the distance of one of the proof masses with respect to the optical bench,

- 3) and 4) two interferometers to assess the residual frequency noise of the laser.

The optical bench is made from low-expansion glass (Zerodur), the optical components are attached using hydroxy-catalysis bonding, a technique developed for the GP-B (Gravity Probe-B) mission to ensure the long-term stability of the components' position.

Once in orbit the residual differential acceleration noise of the proof masses is measured. In order to be able to measure differential acceleration, the sensitive axes of the two test-masses have to be aligned. This requires the development of a capacitive suspension scheme that carries one or both test-masses along with the spacecraft, including along the measurement axis. In LISA Pathfinder (LPF), the optical metrology system essentially makes two measurements; the separation of the test masses, and the position of one test mass with respect to the optics bench. The latter measurement is identical to the LISA local measurement interferometer, thereby providing an in-flight demonstration of precision laser metrology directly applicable to LISA. - Hence, this minimal instrument concept of LTP on LISA Pathfinder is deemed to contain the essence of the construction procedure needed for LISA and thus to demonstrate its feasibility.


Figure 14: Schematic of the basic LTP metrology concept (image credit: ESA)

From a control point of view, the major task of the drag-free system is the stabilization of the test mass relative coordinates (12 DoF) and the spacecraft attitude (3 DoF). In total these are 15 DoF to be controlled. In the ”science” mode, this must be accomplished while minimizing any non-gravitational acceleration along the ”sensitive axis”. See Figure 13.

In LISA and in LPF (LISA Pathfinder), charging by cosmic rays is a major source of disturbance, thereby each test-mass carries a non contacting charge measurement and neutralization system based on UV photoelectron extraction. An in-flight test of this device is then obviously a key element of the overall LPF test.

Each proof (test) mass is surrounded by a set of electrodes that are used to read out the mass position and orientation relative to the spacecraft. This measurement is obtained as the motion of the proof mass varies the capacitances between the electrodes and the proof mass itself. The same set of electrodes is also used to apply electrostatic forces to the proof masses. Differential capacitance variations are parametrically read out by a front-end electronics composed of high accuracy differential inductive bridges excited at about 100 kHz, and synchronously detected via a phase sensitive detector.

Each proof mass, with its own electrode housing, is enclosed in a high vacuum chamber which is pumped down to 10-5 Pa by a set of getter pumps. The laser interferometer light passes through the vacuum chamber wall through an optical window. The free-falling system formed by one test mass, its electrode housing, the vacuum enclosure and the other subsystems is referred to as the GRS (Gravity Reference Sensor).

Within the LTP, a key element for suppressing the force disturbance is that the proof masses have no mechanical contact to the spacecraft. In addition, as forces may depend on the position of the proof masses within the spacecraft, this is kept as fixed as possible.

To fulfil both of these apparently conflicting requirements, the spacecraft actively follows the proof (or test) mass located within it in a closed-loop control scheme referred to as drag-free control. The position of the proof mass relative to some nominal origin is measured by means of the gravitational reference sensor. A high gain control loop tries to null this error signal by forcing the spacecraft to follow the proof mass. In order to produce the necessary force on the spacecraft, the control loop drives a set of micro-thrusters. 36)


Figure 15: Forces and torques acting on the spacecraft and test mass (image credit: ESA)


Figure 16: Conceptual view of the drag-free control loop (image credit: ESA)


Figure 17: Functional units of the drag-free system (image credit: EADS Astrium GmbH)

The interferometer system provides the following measurements:

1) Heterodyne measurement of the relative position of the proof masses along the sensitive axis

2) Heterodyne measurement of the position of one of the proof masses (proof mass 1) relative to the optical bench

3) Differential wave-front sensing of the relative orientations of the proof masses around the y-axis and the z-axis

4) Differential wave-front sensing of the orientation of proof mass 1 around the y-axis and the z-axis. Sensitivities at mHz (milli Hertz) frequency are in the range of 10 pm (1 picometer = 10-12 m) Hz-1/2 for displacement and of 10 nrad Hz-1/2 in rotation.

A diode-pumped, monolithic Nd:YAG non-planar ring oscillator (wavelength 1.064 µm) is used as the light source for the heterodyne interferometry. To obtain the necessary frequency shift, the beam coming from the laser is split and each partial beam is sent through an AOMU (Acousto-Optical Modulator Unit). The light is then delivered to the optical bench by a pair of optical fibers and fiber injectors. Quadrant photo-diodes are used for the detection of the interferometric signal, permitting the measurement of yaw and pitch of the proof masses with respect to the sensitive axis.


Figure 18: Artist's view of the LTP in January 2008 (image credit: ESA, Ref. 25)

Legend to Figure 18: The partly transparent view reveals the two proof masses: 46 mm large cubes of a gold/platinum alloy, housed in individual vacuum cans. The cubes serve both as mirrors for the laser interferometer (red lightpaths) and as inertial references for the drag-free control system of the spacecraft.

The LTP subsystems are:

• ISS (Inertial Sensor Subsystem)

• CMA (Caging Mechanism Assembly)

• CMD (Charge Management Device)

• FEE (Front End Electronics)

• OBI (Optical Bench Interferometer)

• RLU (Reference Laser Unit)

• AOMU (Acoustic Optic Modulator Unit)

• FEEP (Field Effect Electric Propulsion) with µN thrusters

• DDS (Data and Diagnostic Subsystem)

• OBC (On-Board Computer)

• DFACS (Drag-Free and Attitude Control Subsystem)

The LTP instrument package has a mass of about 125 kg, power consumption of about 150 W, and a size of: 64 cm x 38 cm x 38 cm.


Figure 19: Photo of the LPT during the acoustic tests at ESTEC in September 2008 (image credit: ESA)

Legend to Figure 19: This photograph shows two dummies of the two vacuum chambers which will contain the proof masses in the electrodes housing boxes, and the connecting optical bench.


Figure 20: Photo of the electrode housing box in November 2008 (image credit: ESA)

Legend to Figure 20: Individual elements of the electrodes housing (EH) box. When assembled the EH will house one proof mass within the LTP (LISA Technology Package) on LISA Pathfinder.


Figure 21: Photo of the electrode housing and proof masses in November 2008 (image credit: ESA)


Figure 22: Engineering model of the diode pumped Nd:YAG laser of RLU (image credit: ESA, Tesat)


Figure 23: Photo of the Phasemeter assembly (image credit: University of Birmingham, ESA)


Figure 24: Engineering model of OBI (Optical Bench Interferometer), image credit: EADS Astrium, ESA


LTP instrument overview:

The LTP (LISA Technology Package) instrument consists of two main functional subsystems, namely the ISS (Inertial Sensor Subsystem), also known as ISH (Inertial Sensor Head), and the OMS (Optical Metrology Subsystem) as shown in Figure 27, which are controlled by the DDS ( Data & Diagnostics Subsystem).

• The ISH is providing all technical means necessary to bring the two LTP TM (Test Masses) into orbit and then - steered by DFACS algorithms - to control the TM attitude and position through electrostatic actuation and suspension so that one TM is freely falling in direction towards the other. The ISS comprises six main subsystems (Ref. 17):

- Test mass (a glod-platinum test mass)

- Electrode housing

- Two GPRM (Grabbing, Positioning and Release Mechanisms)

- A launch lock, incorporating a venting gate valve, called the CVM (Caging and Venting Mechanism) 37)

- UV discharge system (fiber feedthroughs)

- Vacuum System

- Front end electronics


Figure 25: Photos of the ISH hardware (image credit: ESA)

There are two ISH (Inertial Sensor Heads) on either end of the LTP (LISA Technology Package), mounted to a Zerodur support structure, to which the optical bench is also fixed (Figure 26). Each ISH carries a test mass, which will be free floating once the spacecraft is on orbit. Each of these test masses will be floating inside a set of parallel electrodes, which measure the distance from the mass to its walls. The distance between the two test masses will be measured using a laser interferometer, to confirm that the masses can be maintained in a genuinely free-floating condition inside the spacecraft and that the distance between them can be measured with sufficient accuracy and noise level. These elements will be vital for any future gravitational wave detection mission, which will need to place the masses much farther apart in space. 38)

Creating the correct environment for the test masses poses several challenges: the structure must be perfectly aligned down to micron level, even when it is subjected to the extreme acoustic and mechanical environment of launch. The vacuum enclosure, which is almost 43cm tall and almost 18cm in diameter, must avoid magnetic components: for this reason it was constructed from titanium instead of the more conventional steel, and special feedthroughs were required for UV fibers and electrical cables. The ISH must be constructed from low outgassing materials, to attain a high vacuum level, and this vacuum level must be preserved up to launch. The Caging and Venting Mechanism, which functions as a launch lock, ensures that any tiny air pockets – "virtual leaks" from within the environment – are vented to space, preserving the vacuum.

The mechanical performance of the ISH during launch was verified by a series of tests. The campaign included vibration testing in three planes, carried out by CGS (Carlo Gavazzi Space) at the Centro Technica in Milan, Italy. After a functional test of the mechanism by which the test mass is handed over from the CSV to the GPRM, the ISH was first subjected to swept-sine vibrations at low level, to determine its characteristics and to reveal any mechanical resonances not predicted by the design models. The ISH was then subjected to random vibrations at qualification level – these tests verify the overall design of the ISH. This was followed by a random vibration life test (three cycles at acceptance level – to verify that the performance satisfies the specifications).

The integration of all the components of the ISH with perfect alignment, and the successful completion of the tests mark a major milestone for LISA Pathfinder. It is the first time that a heavy test mass inertial sensor has been assembled and tested successfully anywhere in the world. All flight model elements of the ISH have been delivered and work continues on the LISA Technology Package, due for completion in November 2014, after which it will be delivered and integrated into the spacecraft.


Figure 26: Schematic diagram of the Inertial Sensor Head (image credit: RUAG & CGS)

• The OMS serves as a high precision optical sensor of the differential movement of the two TM and of the movement of one of the TM with respect to the LPF SCM (Science Module). It is based on heterodyne Mach-Zender interferometry allowing for high precision measurements of TM position and attitude, e.g. intrinsically reaching the range of 6 x10-12 m/Hz1/2 x [1+(f/3 mHz)2] for 3 - 30 mHz in case of position sensing.

OMS comprises four main subsystems:

- Reference Laser Unit

- Acousto-Optic Modulator

- Optical Bench

- Phasemeter.


Figure 27: Functional elements of the LTP instrument package (image credit: EADS-Astrium GmbH) 39)


Figure 28: LTP core assembly configuration with the two vacuum enclosures for the test masses (image credit: EADS-Astrium) 40)


Figure 29: LTP accommodation in the LPF spacecraft center (image credit: EADS-Astrium GmbH)

Data acquisition, conditioning and phase measurement is performed by the interferometer front-end electronics, based largely on field programmable gate arrays (FPGA). The final processing and retrieval of the position signals from the phase measurements is performed by the LTP payload computer. 41)


FEEP (Field Effect Electric Propulsion) Subsystem:

The LISA Pathfinder MPS (Micro-thrust Propulsion Subsystem) is comprised of three main subsections called MPA (Micro-Propulsion Assemblies), each one consisting of one FCA (FEEP Cluster Assembly), one PCU (Power Control Unit), and one NA (Neutralizer Assembly) as shown in Figure 30. The main features of the propulsion system are: 42) 43) 44) 45) 46)

• It is able to produce stable thrust levels ranging from 0.1 µN to 150 µN

• It is able to produce thrust with a resolution capability better than 0.1 µN and time response better than 200 ms for any specified thrust step in the required thrust range

• Thrust noise, as measured indirectly through electrical parameters and through direct beam sampling, is compatible with the requirement for proper DFACS operation

• Once deployed and initialized in orbit, it has no moving parts, nor gas leaks that could result in spacecraft disturbance

• The thruster does not need ferromagnetic materials, and the magnetic disturbance on the test mass can be prevented by adequate design rules.

Each MPA is mounted on the spacecraft at 120º with respect to the others with each FCA allocating four FEEP thrusters. The four FEEP thrusters, commanded individually and working in hot redundancy, are mounted in such a way that the relevant thrust vector directions allows to maximize the spacecraft positioning in all spacecraft directions (each thrust vector has 45º of nominal azimuth angle and 30º of nominal elevation). The PCU is allocated inside the spacecraft while the FCA and NA (Neutralizer Assembly) are mounted externally.

The NA (Neutralizer Assembly) consists of a self-contained unit of two Neutralizer unit mounted on a support structure with any necessary interfaces and support bracket (mechanical, thermal and electrical). The neutralizer is necessary to nullify the spacecraft unbalance due to ion thruster operation. The neutralization function is implemented by means of cold redundant hardware.

The PCU (Power Control Unit) consists of a self-contained electronic unit mounted on a support structure with any necessary interfaces and support bracket (mechanical, thermal and electrical). The PCU interfaces the spacecraft (Power and TC/TM tasks) and provides power and control to both FEEP Cluster and Neutralizer assemblies. The HV interconnection box and relevant harness is part of this equipment.


Figure 30: Spacecraft and MPS (Micro-Propulsion Subsystem) layout (image credit: ESA)

The LISA Pathfinder FEEP Subsystem has been developed to embark two different FEEP thrusters technologies currently under qualification in Europe: one using slit-shaped emitter with Cesium as propellant and the second using a needle-shaped emitter with Indium as propellant.


Figure 31: Photos of the needle FEEP (left) and slit FEEP (right) FCAs (FEEP Thruster Assemblies), image credit: ESA

• In the slit-shaped emitter design (developed by Alta SpA, Italy), field emission is generated applying the intense electric field on the liquid metal (Cesium), which is heated above its melting point (≈ 29ºC), inside the edge of two sharp blades so forming the emitter slit. The equilibrium between the surface tension and the electric field strength forms the so-called Taylor cone on the surface with a jet protruding due to space charge. Atoms are then ionized at the tip of the jet and accelerated out by the same field that created them. This configuration allows to form several field emission sites (Taylor cones) along the length of the slit directly proportional to the commanded thrust. With this design approach the thrust range can be extended simply increasing the length of the slit. 47) 48)

The FT-150 FEEP (Field Emission Electric Propulsion) microthruster is designed for extremely fine positioning and attitude control applications. It generates thrust by ejecting cesium ions at about 100 km/s of speed with a noise level lower than the threshold of the nano-balance used in the direct measurement, about 0.1 µN/√Hz, in the 10 mHz to 10 Hz range. Ions are extracted from the emitter tip and accelerated by the electric field created between the emitter and the accelerator electrode placed in front of it. The total voltage applied to the electrodes is between 7 kV and 13 kV. The specific impulse of the FT-150 FEEP microthruster is substantially larger than the typical range of ion thrusters, varying between 3000 s and 4500 s depending on operating conditions. Performance was verified by testing throughout a thrust range from 0.1 µN to 150 µN. Alta SpA (Italy) carried out the development of the FT-150 FEEP microthruster with the aim to fulfill the requirements of the fundamental physics missions LISA and Microscope, as well as those of the LISA technology demonstrator LISA Pathfinder. 49)




Nominal power

6 W

@ 100 µN of thrust

Thrust range

0.1 to 150 µN


Thrust resolution

Below 100 nN


Thrust accuracy

± 1.6 µN at max thrust
± 0.4 µN up to 4 µN


Thrust response time

50-150 ms


Thrust noise

< 0.1 µN/√Hz

Below nano-balance detection threshold

Specific impulse

About 6000 s (depending on emitter voltage)


Total impulse capability

> 5000 Ns

2631 Ns demonstrated by endurance test

Thruster dry mass

~1400 g


Propellant mass

92 g

Per thruster

Table 4: FT-150 FEEP microthruster performance data

• In the needle-shaped design [developed by ARC (Austrian Research Centers GmbH) of Schreibersdorf, Austria], field emission is generated applying the intense electric field on the liquid metal (Indium), which is heated above its melting point (≈ 156ºC), on a needle-shaped configuration. In this case, due to limited thrust capability of a single needle (it allows a single site emission only), a cluster of several needles need to be used to form a thruster with suitable thrust range (e.g. nine to cover LISA Pathfinder requirements). The indium happens to be less reactive than cesium, hence it can be more easily managed during all phases of AIT and mission. 50) 51) 52)

The two techniques are equivalent as far as the propulsion performances are concerned; however, the slit cesium design requires less electrical power to liquefy the propellant with respect to needle-indium due to lower melting temperature (29ºC against 156ºC). The mass budget is similar for thrust values of about 100 µN and is resulting favorable for the needle in case of lower thrust range and, conversely, beneficial for slit in case of higher thrust range.


Figure 32: FEEP thruster element (image credit: ESA)

Legend to Figure 32: On the left side the extractor electrode [2], the focusing electrode [3] and the cover-plate [4]. On the right side the FEE.

NA and neutralizer: For LISA Pathfinder the neutralizer has been configured as a self-standing unit in a dedicated mechanical box where two neutralizers (main and redundant) have been allocated. The so formed NA (Neutralizer Assembly) is allocated on the external wall of the satellite and not on the same panel where the FCA is located. As a consequence of this choice and taking into account of the LISA Pathfinder space environment (L1 with very low plasma density), the neutralization function is assured by means of additional bias voltage (200V max) to enhance electron emission. Consequently being allocated far from FCA that has reduced effects of propellant contamination and electric field caused by the FEEP operation. - The neutralizer equipments for LISA Pathfinder has been developed by Thales-Alenia Space, Florence, Italy (TAS-I). 53)


Figure 33: The EQM (Engineering Qualification Module) of the neutralizer assembly (image credit: (TAS-I)

PCU (Power Control Unit): The requirements of the PCU are to control any thrust in the range of 0.1 to 150 µN with a resolution better than 0.1 µN. Such a type of PCU has been developed by Galileo Avionica whose architecture can be adapted with different electrical and mechanical arrangements, providing the following main features: 54) 55)

• Control and management up to four independent FEEP thrusters working in hot redundancy, providing operating voltage up to 13.7 kV at very low currents (from 0.5 µA to 2 mA)

• Control and management of two neutralizers working in cold or hot redundancy

• Single point failure tolerant architecture for its use as “primary” propulsion allowing at least three FEEP thrusters and one neutralizer fully operating in case of single failure at FEEP subsystem level (fully redundancy concept also applicable for command, telemetry and Bus power interfaces).

The most important task is to control the thrust level and to this purpose the PCU embarks two dedicated HV (High Voltage) supplies for each thruster with the main characteristics provided in Table 5. These high voltage supplies are tailored according to specific thruster needs within a maximum total voltage range of 13.7 kV. Voltages, currents and telemetries are suitable to match the two different FEEP thrusters technologies: slit-shaped emitter with Cesium propellant and needle-shaped emitter with Indium propellant.


PCU capability

Emitter voltage (Ve)

From 0 to 12 kV

Emitter current (Ie)

From 0 to 2 mA

Emitter power

up to 18 W

Accelerator voltage (Va)

From -1 to -1.7 kV

Accelerator current (Ia)

From 0 to 1 mA

Accelerator power

up to 1.7 W

TM monitoring

emitter voltage, beam current (Ie – Ia), accelerator voltage, accelerator current, arc discharge counter

Table 5: PCU HV (High Voltage) supply capability

For thrust regulation, the implemented closed loop control is allocated in the PCU and follows the control principle shown in Figure 34. It is composed of one inner analogue current loop tracking firstly the beam current parameter and then an overall digital control loop with 12 bit of resolution tracking both emitter voltage and beam current.


Figure 34: Schematic of the thrust control loop (image credit: SELEX Galileo)

The thrust control performances (from the electrical point of view) have been verified with both FEEP thrusters technologies (slit and needle) providing performances in line with the requirements. The qualification of the PCU for LISA Pathfinder has been complemented by a dedicated qualification and endurance test of the high voltage power boards, exploiting the HV functions, to cover with suitable safety margin the emitter voltage extreme (i.e. in excess of 15 kV and 20% of additional power level).

On LISA Pathfinder, the capability of adaptation to different thrusters technology and neutraliser solution, without jeopardising the PCU design and implementation, reveals an excellent level of flexibility, and adequate for potential use of this PCU for future missions.


Figure 35: Photo of the PCU flight model for Lisa Pathfinder (image credit: SELEX Galileo)


DRS (Disturbance Reduction System):

DRS is a NASA-provided instrument package within the New Millennium Program developed by NASA/JPL. When first proposed, the DRS payload was known as DRS-PFCV (Disturbance Reduction System-Precision Flight Control Validation), consisting of the GRS (Gravity Reference Sensor) design of Stanford University, which closely resembled the LTP, namely in that it consisted of two inertial sensors with the associated interferometric readout, as well as the drag-free control laws and µN colloidal thrusters (organic ionic liquid) - although the technologies employed were different from the LTP implementation. 56) 57) 58) 59) 60) 61) 62)

Due to budgetary problems, the descoped DRS (Oct. 2005) of the former ST7 (Space Technology 7) mission now consists of the micro-Newton (µN) colloidal thrusters, DFACS (Drag-Free and Attitude Control System), and a microprocessor. The DRS will now use the LTP inertial sensors as its drag-free sensors (test masses position and attitude) to control the spacecraft attitude with independent drag-free software and will use the colloidal thrusters as actuators.

Note: There are, in fact, two separate “DFACS packages” integrated on board the LISA Pathfinder spacecraft. LTP will utilize its own DFACS algorithm during its allocated operations period, which occurs before the DRS operations period. - During DRS operations, NASA will use its own IAU (Integrated Avionics Unit) package which includes DCS (Dynamics Control Software).

The primary goal of the DRS instrument package is to maintain the position of the spacecraft with respect to the proof mass of LTP to within 10 nm Hz-1/2 over the frequency range of 1-30 mHz. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass.

The conceptual functionality of the DRS system is shown in Figure 36. The two cubical test masses TM1 and TM2 are enclosed within housings rigidly attached to the body of the spacecraft. Electrodes on the inner faces of the housings are used to measure the position and orientation of the test masses with respect to the housings using a capacitive sensing mechanism. A laser interferometer is being used to measure the distance changes between the two test masses to infer the residual acceleration noise. Colloidal microthrusters are being used to counteract the external forces, which are primarily due to solar radiation pressure acting on the spacecraft solar panel. The thrust level is continually being adjusted to keep the spacecraft centered about the test masses.


Figure 36: Conceptual diagram of the DRS system (image credit: NASA/JPL)

Disturbances: The largest disturbances to the inertial trajectory of a spacecraft (radiation pressure, residual gas drag, and particulate impacts) are cancelled by the basic concept of a drag-reduction system. The final performance of the system will be limited by a number of smaller disturbances. These disturbances fall into three categories:

1) Variations in the gravitational potential at the test-mass location

2) Momentum transfer to the test mass by residual gas and cosmic radiation particles

3) Variations of the electromagnetic fields at the test-mass location.

The main gravitational fluctuations are due to the thermal distortion of the spacecraft and to the relative displacement of the test mass with respect to the spacecraft. Reducing the gravity gradient and displacement of the test mass minimizes the gravity noise caused by spacecraft displacement.



Thrust range

5 to 30 µN

Thrust precision

< 0.1 µN

Thrust noise

< 0.1 µN Hz-1/2 (5 Hz control loop)

Thrust command rate

10 Hz (< 0.1 s latency)

Thrust respond time

< 100 s from maximum to minimum

Specific impulse (30 µN point)

> 150 s

Specific impulse (6 µN point)

> 275 s

Operational lifetime

> 2,200 hours

Plume half angle

< 35% (95% beam current)

Table 6: Summary of the DRS microthrust propulsion system requirements

CMNTA (Colloid Micro-Newton Thruster Assembly), designed and developed for NASA by Busek Co. Inc. of Natick, MA. The objective is to smoothly and continuously counter all external disturbances with control authority over all six degrees of freedom (DoF) of the spacecraft motion. DRS provides a thrust level range of 5-30 µN with a resolution of 0.1 µN.
Note: Colloid thruster types are part of the EIP (Electrostatic Ion Propulsion) family. The EIP concept uses a high voltage electrostatic field to accelerate positively charged particles (or ions) to large exhaust velocities (acceleration is created by the force on charged particles in the electric field).

The CMNTA thrusters use a colloidal fluid propellant. The fluid is fed through a needle by a pressurizing system. At the tip of the needle, a high electrical field is applied, which causes droplets to form and to be ejected from the tip of the needle. The droplets of the “electrospray” are spontaneously charged and accelerated by the electric field. A typical single-emitter-needle thruster produces a maximum thrust of 3 µN. Each thruster employs an array nine needles while four thrusters are mounted on one “cluster” assembly. Two clusters of thrusters are being used for DRS. Each cluster consists of the thrusters, one carbon nanotube emitter, the propellant feed system, and the PPU (Power Processing Unit). The PPU contains all the DC-DC converters to power the system and the autonomous controls for the carbon nanotube field emission neutralizer. The full thruster cluster is shown in Figure 37.

The CMNT key elements are: 63) 64)

• Thruster head: The thruster head is comprised of a manifold that feeds nine emitters and the electrodes that extract and accelerate the propellant.

• Propellent feed system: Propellant is stored in a stainless steel bellows compressed by four constant force springs set to supply the microvalve with propellant at approximately 1 atmosphere of pressure. A µValve is piezo-actuated using ~1 mW of power to control the propellant flow rate to better than 1 nA equivalent resolution. This level of precision corresponds to ≤ 0.01 N of thrust, with a response time over its full range of less than 0.5 s.

• Cathode neutralizer: The cathode neutralizer is made from a carbon nanotube (CNT) base with an extractor electrode. The cathode is capable of producing 10 µA to 1 mA using extraction voltages of 250-770 V.

• Thruster electronics: The thruster electronics consists of 4 power processing units (PPUs) and one digital control and interface unit (DCIU) for each cluster. A PPU includes the high-voltage DC-DC converters.


Figure 37: Photographs of the DRS flight hardware, showing the two clusters of µN Colloidal thrusters, and the IAU (image credit: NASA/JPL, Busek, Ref. 12)

IAU (Integrated Avionics Unit), designed and built for NASA by Broad Reach Engineering Inc., Tempe, AZ. The flight software resides in IAU which serves as the interface among the drag-free sensors, the thrusters, and the host spacecraft. [Note: In ESA terminology, the IAU system is referred to as DFACS (Drag-Free and Attitude Control System) + microprocessor].

The IAU (Figure 38) contains a 30 MHz cPCI (Compact Peripheral Component Interface) backplane, a Rad-750 processor (main processing board), and specialized cards [CAPI (Command and Payload Interface), SMACI (State Monitor and Attitude Control Interface)] to support communications with the spacecraft and the thrusters, as well as housekeeping sensor monitoring (temperature and currents). The IAU relays science data to the spacecraft, which is responsible for downlinking data to the ground.

The DCS (Dynamics Control Software) is also part of the flight software. The flight software executes the DCS at 10 Hz. The spacecraft interface provides position and attitude measurements from the drag-free sensors, as well as the attitude and rates of the spacecraft. The DRS sends requested test mass forces and torques to the drag-free sensors. This supplements the force and torque commands sent to the colloid thrusters to act on the spacecraft.

The DCS determines the thruster commands to control the spacecraft position and attitude based on the measurements of the position of each test mass relative to its housing. The variation in thrust commanded by DCS must be within the response capability of the thrusters. The electrostatic forces and torques for the test masses are a function of the test-mass housings. The spacecraft control requirement is to keep the spacecraft centered about the two test masses within < 10 nm Hz-1/2 over the frequency range of 1 to 30 mHz.

Several disturbance models are included in the design of the controls: solar radiation pressure variation; capacitive sensing noise (modeled as a colored power spectrum); thruster and star tracker noise (modeled as white); and acceleration noise on the test mass, including magnetic and Lorentz forces, thermal variations (self gravity), and cosmic ray impacts.


Figure 38: Block diagram of the IAU (image credit: NASA/JPL)


Figure 39: CMNT cluster functional block diagram with pictures of various components (image credit: NASA/JPL, Busek)


Figure 40: Layout of the micro-propulsion subsystem (image credit: ESA)



Ground segment:

The LISA Pathfinder ground segment comprises two operational centers, both provided by ESA:

• The MOC (Mission Operations Center) at ESA/ESOC. MOC is responsible for LEOP, the transfer phase, and all operations during the routine phase and is in contact with the spacecraft for eight hours per day through the ground station(s).

STOC (Science and Technology Operations Center) is the point of interface to the scientific community, and is responsible for the payload scheduling (both long and short-term), quick-look data analysis, data processing and archiving.

The STOC will also take a leading role in the analysis of the mission data. Development of the STOC is run from the ESAC (European Space Astronomy Center) in Villafranca, Spain, however, during the science operations of the LTP, the STOC will be re-located to ESOC. This is to enable the required close contact between the science operations planing and the mission operations (Ref. 5). 65)


Figure 41: Overview of the LISA Pathfinder ground segment (image credit: ESA)

The main activities of the STOC fall into the following classes:

• LTPP (Long-Term Payload Planning): This activity is concerned with the high-level planning of the 90 days of the mission operations. In particular, in defining the experiments to be performed, and the creation of a strawman operational plan. The results of the LTPP are contained in the EMP (Experiment Master Plan)

• MTPP (Medium-Term Payload Planning): This activity concerns the validation of the POR (Payload Operational Requests). A POR is a time-tagged list of telecommand sequences to be executed autonomously. One POR is required for each day of operations.

• STPP (Short-Term Payload Planning): This activity deals with the delivery of validated PORs to the MOC for generation of the Mission Timeline.

• DI (Data Ingestion): The main purpose of the STOC Data Ingestion System is to retrieve telemetry data during each of the ground passes and to make it available to the STOC Quick-Look and DA (Data Analysis) and to the archive system for future usage by LTP and DRS.

• QL (Quick Look): The aim of the QL subsystem is to monitor the LTP operations taking place and to provide an alert in case something is not as expected. The QL uses a subsystem of full science telemetry which is directed to a specific packet-store board, and is telemetered with high priority at the start of the ground station pass. Following the QL activities, the STOC may:

- Issue a warning for a deeper investigation as part of the DA.

- Request to MOC a change of TC parameter to be applied to the next run if an immediate action is needed.

- Request to MOC to immediately command the LTP into Standby mode.

• DA (Data Analysis): The DA is a joint effort of the STOC, LTP and DRS teams and will use the telemetry and auxiliary files available in the STOC archive.

• SA (Science Archive): The SA will make all the data accumulated by LPF and a subset of the data analysis products available to the wider scientific community. The LTP team have priority rights to the data for the first three months, after which the archive goes live on the public domain.


Figure 42: Overview of the science operations ground segment showing the data flow between the various subsystems (image credit: ESA)

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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.