# Towards NGGM: Laser Tracking Instrument for the Next Generation of Gravity Missions
## 2 LTI Design Principle
The optical setup on the optical bench(es) routed the beam from the fiber collimators via a steering mirror (FSM) to a beam splitter (BS), which split the beam into one large part that was sent to the other spacecraft (via the retroreflector) and one smaller part that was guided to the photoreceivers to act as a local oscillator. The imaging optics in front of the photoreceivers simultaneously imaged the steering mirror and the entrance aperture onto the photoreceivers, thereby minimizing the effect of beam walk due to beam angle changes, as well as phase errors due to diffraction effects of the baffles and entrance aperture. The compensation plate (CP) minimized the ranging noise introduced by the beam splitter under pointing noise. Only the beam splitter and compensation plate were in the direct  
Figure 2: Concept sketch of the laser tracking instrument in off-axis transponder configuration: (**a**) with partial redundancy; (**b**) with full redundancy. measurement path, in which any path length noise (thermally or pointing-driven) directly coupled into the ranging performance. The pathlength noise of all other optical elements (from fibers to beam splitter, and beam splitter to photoreceivers) was strongly suppressed due to common-mode effects in the local oscillator and roundtrip beams. The received beam from the other spacecraft with a power between 0.1 and 3 nW (derived from optical link budget calculations for high and low cases in the optical power budget [19]) entered the optical bench, was reflected at the beam splitter and imaged onto the photoreceivers. The quadrant photoreceivers on the optical bench were operated in hot redundancy. The heterodyne signal (the beat between the local oscillator and the received beam from the other spacecraft) was read out from the four elements of the photoreceivers, pre-amplified and processed in the phasemeter, which was one part of the instrument control units. The phase variations represented the relative distance change of the two spacecrafts relative to each other, which was the main science signal. The phase differences between the individual segments delivered the pointing information to the other spacecraft with respect to the line of sight, used to drive the attitude control system of the satellite to \(\upmu\)rad accuracy using differential wavefront sensing (DWS). The off-axis retroreflector with its vertex in the center of mass (CoM) of the spacecraft routed the beam around the CoM, thereby enabling the distance measurement from CoM to CoM of the two spacecrafts. It needed to provide a beam co-alignment accuracy (incoming to outgoing beam) of less than 40 \(\upmu\)rad and a low-temperature dependency of the vertex position to enable the required ranging performance. At the second spacecraft (with the laser in transponder mode), the identical instrument configuration was implemented, but differing on the two spacecrafts with respect to the orientation and flight direction (which was the same for both spacecrafts). The received signal from the first spacecraft was used to offset lock the local laser by some MHz, and also to point the second spacecraft to the first spacecraft (by use of the DWS signal). Apart from the mode of operation of the laser, the operation principle was the same as on spacecraft one. In addition to the higher reliability [19], the fully redundant LTI allowed for the use of two identical satellites without having to fly one satellite oriented backwards.