Navigation, Guidance & Control Specification Sheet

Kenneth M Rock
August 1995

The Navigation, Guidance & Control (NGC) subsystem onboard the spacecraft will be able to: (1) control the spacecraft state vector (motion of center of mass, it's trajectory) and (2) control the orientation of the spacecraft (motion about center of mass, it's attitude)

Navigation: Determining position and attitude through sensors.
Guidance: Computer processing of navigational information used to maintain trajectory within nominal profile and maintain desired attitude.
Control: Implements guidance decisions for corrections in trajectory or attitude.
The responsibilities of the NGC group include: spacecraft trajectory design, attitude control system design, spacecraft stability, attitude/trajectory sensor selection, attitude effector design requirements, landing trajectory design, and landing site selection.

Trajectory Design & Control


The NGC group will have information regarding the launch vehicle performance based on the vehicle selection of the Propulsion group. The launch vehicle final stage trajectory could be altered by the NGC group for optimum use of the final stage DV.

NGC's first responsibility is to determine a trajectory for the spacecraft, and the required DV's, including a 10 percent error budget. These DV's will be provided to the Propulsion group for Main Propulsion System (MPS) selection. Preliminary trajectory design yielded the DV's tabulated in table 1.0. These values assume a Delta-II launch and a PAM-D upper stage (~1.45 km/sec). Furthermore, the earth orbit plane change maneuver may not be required, depending on launch time selection.

Mission PHASEDelta-V's
a. Earth Orbit Plane Change Maneuver 0.69 km/sec
b. Trans-Lunar Injection 1.67 km/sec
c. Lunar Orbit Insertion 0.54 km/sec
d. De-Orbit Burn 0.012 km/sec
e. Lander Deceleration (to approx. 1000m) 1.7 km/sec
f. Landing Site Selection (approx. 200m x approx. 200m area) unknown
g. Final Lander Descent to MPS Cutoff unknown

Table 1.0

The overall trajectory should include a lunar polar orbit to allow access to all plausible landing sites on the nearside of the moon without requiring direct landing from the translunar orbit. Line of site communication with the DSN shall be possible throughout the entire landing phase.

The during the landing trajectory, the spacecraft must be able to actively examine the predicted landing site for surface inclination, craters, and rubble. Based on processing of sensor data, the spacecraft must decide whether it is safe to land at the predicted site. If it is not, the lander will alter it's descent profile towards a safer landing site. Preliminary estimates defined landing hazards as surface inclination greater than 10, craters larger than 1 meter, and rubble larger than 0.25 meter. These estimates are related to the structural design of the spacecraft.

The NGC group must determine whether the flight trajectory will be controlled predominantly by the spacecraft itself, or by mission control. The preliminary spacecraft design utilized a system which could detect accelerations well enough to calculate its own trajectory. This data could be compared to nominal trajectory for autonomous midcourse corrections, however, complicates the design. Another possibility is to determine the spacecraft trajectory through tracking via the DSN, and uplink data to the spacecraft. This trade could open tolerances to allow selection of alternate components for various reasons which could include higher mission safety.


The NGC group must select sensors to measure the accelerations of the spacecraft. The precision of these instruments depends upon the required accuracy for the navigation that the spacecraft performs onboard (as opposed to ground control). Preliminary design of the WORLD-M spacecraft utilized an Delco 3-Axis Space SIRU-3 Inertial Reference Unit (SIRU-3) for the measurement of the trajectory and attitude. This IMU was originally chosen for an orbiter which was to map the surface of the moon for 1 year. Therefore, for use in the current proposal, a smaller, less precise instrument could possibly be used.

Several sensors will be required for the landing phase. An altimeter is required to accurately measure altitude and rate of descent. Additionally, some type of sensor will be needed to resolve landing site hazards, such as surface inclination, craters, and rubble. This sensor could also serve the purpose as the primary or backup altimeter. The sensor group may be of assistance in choosing these instruments.


The NGC group should determine the computer requirements for trajectory calculations and maneuver planning, and provide this information to the computer group. The NGC group will also provide the necessary information to allow the computer group to integrate the trajectory sensors with the computer system, and then the computer system to the trajectory effectors. Since the effectors will likely be the MPS, the NGC group will need information regarding control of the propulsion system, like burn times, gimballing, and physical limitations.


The effectors for the propulsion system will be provided to NGC by the propulsion group. Coordination will be required to iterate trajectories based on real engine performance instead of idealized impulsive DV's.

Attitude Control System


A main objective of the NGC group is to determine the optimum stabilization method for the spacecraft. Spin-Stabilized, 3-axis stabilized, or dual-spin spacecraft designs should be considered. The preliminary design utilized a 3 axis stabilized platform, however, the omission of the orbiter portion of the spacecraft may alleviate the need for this type of stabilization. The spacecraft could be designed for spin stability during transit and burns, while during the landing phase, the spacecraft could be actively stabilized. The selection of the stabilization method will impact thermal and communications systems design and these trades should be considered.

A six-degree of freedom model (x,y,z translation and rotation) shall be created to examine the trajectory and attitude of the spacecraft with time during the transit to the lunar surface. This model will provide the basis for development of the attitude control system requirements with regard to burn attitudes, landing stability, communications constraints, and solar panel alignment.


Sensors must be chosen to initially determine and to re-acquire the attitude of the spacecraft during any phase of flight. Additional sensors must be chosen which can track the attitude of the spacecraft with time. The required accuracy of measurement instruments are to be determined by the NGC group.

The preliminary WORLD-M design utilized a pair of Clementine star trackers for attitude acquisition. It is possible that sun-sensors and horizon sensors could be used in lieu of star trackers. The preliminary design for attitude measurement utilized an Inertial Measurement Unit. The original IMU was selected to support a moon orbiter. Since that portion of the mission has been eliminated, a new IMU which is less precise, lighter, and more compact could be chosen.


The NGC group should examine the computer requirements for attitude and trajectory maneuver planning (based of 6 DOF model) and provide this information to the computer group. This information will also include data storage, throughput, and code lengths. The NGC group will also provide the necessary information to allow the computer group to integrate the attitude sensors with the computer system, and then the computer system to the attitude effectors. The RCS system integration will require collaboration with the propulsion and structures groups.

The NGC group must select the type of feedback control to be employed by the control system. The preliminary method utilized PID control. The gain selection and system response should be modeled. This information can then be provided to the structures group for dynamic loading, and will provide a basis for the actual onboard control system coding.


The attitude control effectors chosen in the preliminary WORLD-M design were momentum wheels for fine attitude control, and a Reaction Control System (RCS) for coarse attitude control and momentum dumping. The NGC group is required to select components for the fine control system and determine the DV budget and pulse requirements for the RCS.

Fine control requirements are primarily driven by communications antenna pointing requirements, solar panel alignment, and DV maneuver alignment. Preliminary design iteration indicated that the best method of fine attitude control was through the use of momentum wheels. Sizing of the wheels is driven by the rotational inertias of the structure and the time compliance for the maneuver.

The coarse control effectors utilized in preliminary designs is a Reaction Control System (RCS). The RCS is used for large attitude corrections which must be accomplished quickly, for dumping the excess momentum of the fine control system, and to provide authoritative landing phase attitude control. The RCS may further be used for minor DV corrections of the trajectory. The NGC group will provide RCS torque, DV, and impulse requirements to the propulsion group, who will in turn provide real thruster performance for integration into NGC models.

NGC and Subsystem Interactions

The interaction of all NGC subsystem components with other design groups shall be determined completely. NGC components shall conform to the mass, volume, mounting, and structural loading constraints are defined by the structures group. The interfaces (physical and digital) with the computers group shall be well defined, including throughput, code length, and data storage requirements. The power requirement of the NGC components, including peak and average power draws, will be negotiated with the power systems group. Finally, the thermal requirements of all components will be provided to the thermal group.

Navigation, Guidance & Control Interface Sheet

Input Group Interface/Information Output Group

Trajectory Design & Control

Impulsive dV Trajectory DesignPropulsion
PropulsionReal Trajectory Design
PropulsionLaunch Vehicle Final Stage Orbital Trajectory
PropulsionLanding Trajectory DesignPropulsion

Trajectory Control Computer Interface (guidance)Computer

Trajectory Control System Power RequirementsPower

Trajectory Control System Thermal ConstraintsThermal

Communications Occultation (Farside of Moon)Communication

Solar Occultation (Darkside of Earth/Moon)Power

Attitude Control System
CommunicationsCommunications Antenna Pointing Tolerance
PowerSolar Array Pointing Tolerance
PropulsionReal Thruster Performance
StructuresMoments of Inertia vs. Fuel Consumption

Torque Requirements (thruster couples)Propulsion

Attitude Control Computer Interface (control)Computer

Attitude Control System Power RequirementsPower

Attitude Control System Thermal ConstraintsThermal

Navigation, Guidance & Control Reference Sheet

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