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Chapter Contents

Chapter Glossary

(ADCS)Attitude Determination and Control System
(CoCom)Coordinating Committee for Multilateral Export Controls
(COTS)Commercial-off-the-Shelf
(DOF)Degrees of Freedom
(DSAC)Deep Space Atomic Clock
(DSN)Deep Space Network
(EAR)Export Administration Regulations
(FOGs)Fiber Optic Gyros
(GNC)Guidance, Navigation & Control
(GSO)Geo-stationary Orbit
(USAF)U.S. Air Force
(HCI)Horizon Crossing Indicators
(IMUs)Inertial Measurement Units
(JPL)Jet Propulsion Laboratory
(Lidar)Light Detection and Ranging
(LMRST)Low Mass Radio Science Transponder
(MarCO)Mars Cube One
(PMSM)Permanent-magnet Synchronous Motor
(SDST)Small Deep Space Transponder
(SWaP)Size, weight, and power
(TLE)Two-Line Element
(TRL)Technology Readiness Level

5.1 Introduction

The Guidance, Navigation & Control (GNC) subsystem includes the components used for position determination and the components used by the Attitude Determination and Control System (ADCS). In Earth orbit, onboard position determination can be provided by a Global Positioning System (GPS) receiver. Alternatively, ground-based radar tracking systems can also be used. If onboard knowledge is required, then these radar observations can be uploaded and paired with a suitable propagator. Commonly, the U.S. Air Force (USAF) publishes Two-Line Element sets (TLE) (1), which are paired with a SGP4 propagator (2). In deep space, position determination is performed using the Deep Space Network (DSN) and an onboard radio transponder (3). There are also technologies being developed that use optical detection of celestial bodies such as planets and X-ray pulsars to calculate position data (23).

Using SmallSats in cislunar space and beyond requires a slightly different approach than the GNC subsystem approach in low-Earth orbit. Use of the Earth’s magnetic field, for example, is not possible in these missions, and alternate ADCS designs and methods must be carefully considered. Two communication relay CubeSats (Mars Cube One, MarCO) successfully demonstrated such interplanetary capability during the 2018 Insight mission to Mars (4). This interplanetary mission demonstrated both the capability of this class of spacecraft and the GNC fine pointing design for communication in deep space.

ADCS includes sensors to determine attitude and spin rate, such as star trackers, sun sensors, horizon sensors, magnetometers, and gyros. In addition, the ADCS is often used to control the vehicle during trajectory correction maneuvers and, using accelerometers, to terminate maneuvers when the desired velocity change has been achieved. Actuators are designed to change a spacecraft’s attitude and to impart velocity change during trajectory correction maneuvers. Common spacecraft actuators include magnetic torquers, reaction wheels, and thrusters. There are many attitude determination and

Miniaturization of existing technologies is a continuing trend in small spacecraft GNC. While three-axis stabilized, GPS-equipped, 100 kg class spacecraft have been flown for decades, it has only been in the past few years that such technologies have become available for micro- and nano-class spacecraft. Table 5-1 summarizes the current state-of-the-art of performance for GNC subsystems in small spacecraft. Performance greatly depends on the size of the spacecraft and values will range for nano- to micro-class spacecraft.

The information described below is not intended to be exhaustive but provides an overview of current state-of-the-art technologies and their development status for a particular small spacecraft subsystem. It should be noted that Technology Readiness Level (TRL) designations may vary with changes specific to payload, mission requirements, reliability considerations, and/or the environment in which performance was demonstrated. Readers are highly encouraged to reach out to companies for further information regarding the performance and TRL of described technology. There is no intention of mentioning certain companies and omitting others based on their technologies or relationship with NASA.

Table 5‑1: State-of-the-Art GNC Subsystems
ComponentPerformanceTRL
Reaction Wheels0.00023 – 0.3 Nm peak torque, 0.0005 – 8 N m s storage7-9
Magnetic Torquers0.15 A m2 – 15 A m27-9
Star Trackers8 arcsec pointing knowledge7-9
Sun Sensors0.1° accuracy7-9
Earth Sensors0.25° accuracy7-9
Inertial SensorsGyros: 0.15° h-1 bias stability, 0.02° h-1/2 ARW
Accels: 3 µg bias stability, 0.02 (m s-1)/h-1/2 VRW
7-9
GPS Receivers1.5 m position accuracy7-9
Integrated Units0.002-5° pointing capability7-9
Atomic Clocks10 – 150 Frequency Range (MHz)5-6
Deep Space NavigationBands: X, Ka, S, and UHF7-9
Altimeters~15 meters altitude, ~3 cm accuracy7

5.2 State-of-the-Art – GNC Subsystems

5.2.1 Integrated Units

Integrated units combine multiple different attitude and navigation components to provide a simple, single-component solution to a spacecraft’s GNC requirements. Typical components included are reaction wheels, magnetometer, magnetic torquers, and star trackers. The systems often include processors and software with attitude determination and control capabilities. Table 5-2 describes some of the integrated systems currently available that are associated with a TRL value of 7-9. Blue Canyon Technologies’ XACT (figure 5.1) flew on the NASA-led missions MarCO and ASTERIA, both of which were 6U platforms, and have also flown on 3U missions (MinXSS was deployed from NanoRacks in February 2016).

computer rendering of harware
Figure 5.1: BCT XACT Integrated ADCS Unit.
Credit: Blue Canyon Technologies
Table 5-2: Currently Available Integrated Systems
ManufacturerModelMass (kg)ActuatorsSensorsProcessorPointing Accuracy
ArcsecArcus ADC0.7153 reaction wheels 3 magnetic torquers1 star tracker  3 gyros  6 photodiodes 3 magnetometersYes0.1°
Berlin Space TechnologiesIADCS-1000.43 reaction wheels
3 magnetic torquers
1 star tracker
3 gyros,
1 magnetometer,
1 accelerometer
Yes<<1 deg
AAC Clyde SpaceiADCS-2000.4703 reaction wheels
3 magnetic torquers
1 star tracker
1 IMU, Optionally high precision magnetometer and sun sensors
Yes<1°
AAC Clyde SpaceiADCS-4001.73 reaction wheels
3 magnetorquers
1 star tracker, 1 IMU, Optionally high precision magnetometer and sun sensorsYes<1°
Blue Canyon TechnologiesXACT-150.8853 reaction wheels
3 magnetorquers
1 star tracker
3-axis magnetometer
Yes0.003/0.007°
Blue Canyon TechnologiesXACT-501.2303 reaction wheels
3 magnetorquers
1 star tracker
3-axis magnetometer
Yes0.003/0.007°
Blue Canyon TechnologiesXACT-1001.8133 reaction wheels
3 magnetorquers
1 star tracker
3-axis magnetometer
Yes0.003/0.007°
Blue Canyon TechnologiesFlexcoreconfiguration dependent3 – 4 reaction wheels
3 magnetorquers
2 star trackers
3-axis magnetometer
Yes0.002°
CubeSpace Satellite SystemsCubeADCS 3-Axis Small0.553 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
Yes<1°
CubeSpace Satellite SystemsCubeADCS 3-Axis Small with Star Tracker0.613 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
1 star tracker
Yes<0.1°
CubeSpace Satellite SystemsCubeADCS 3-Axis Medium0.793 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
Yes<1°
CubeSpace Satellite SystemsCubeADCS 3-Axis Medium with Star Tracker0.843 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
1 star tracker
Yes<0.1°
CubeSpace Satellite SystemsCubeADCS 3-Axis Large1.13 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
Yes<1°
CubeSpace Satellite SystemsCubeADCS 3-Axis Large with Star Tracker1.153 reaction wheels
3 magnetorquers
10 coarse sun sensors
2 fine sun/earth sensors
1 magnetometer
1 star tracker
Yes<0.1°
CubeSpace Satellite SystemsCubeADCS Y-Momentum0.31 momentum wheel
3 magnetic torquers
10 coarse sun sensors
1 magnetometer
Yes<5°

5.2.2 Reaction Wheels

Miniaturized reaction wheels provide small spacecraft with a three-axis precision pointing capability. They must be carefully selected based on several factors including the mass of the spacecraft and the required rotation performance rates. Reaction wheels provide torque and momentum storage along the wheel spin axis which results in the spacecraft counter-rotating around the spacecraft center of mass due to conservation of angular momentum from the wheel spin direction. Table 5-3 lists a selection of high-heritage miniature reaction wheels. Except for three units, all the reaction wheels listed have spaceflight heritage. For full three-axis control, a spacecraft requires three wheels mounted orthogonally. However, a four-wheel configuration is often used to provide fault tolerance (6). Reaction wheels need to be periodically desaturated using an actuator that provides an external torque, such as thrusters or magnetic torquers (7).

In addition, the multiple reaction wheels are often assembled in a “skewed” or angled configuration such that there exists a cross-coupling of torques with two or more reaction wheels. While this reduces the torque performance in any single axis, it allows a redundant, albeit reduced, torque capability in more than one axis. The result is that should any single reaction wheel fail, one or more reaction wheels are available as a reduced-capability backup option.

Table 5-3: High Heritage Miniature Reaction Wheels
ManufacturerModelMass (kg)Peak Power (W)Peak Torque (Nm)Momentum Capacity (Nms)#WheelsRadiation Tolerance (krad)
Berlin Space TechnologiesRWA051.7000.50.0160.5130
Blue Canyon TechnologiesRWP0150.13010.0040.0151Unk
Blue Canyon TechnologiesRWP0500.24010.0070.0501Unk
Blue Canyon TechnologiesRWP1000.33010.0070.1001Unk
Blue Canyon TechnologiesRWP5000.75060.0250.5001Unk
Blue Canyon TechnologiesRW10.950100.071.0001Unk
Blue Canyon TechnologiesRW43.200100.2504.0001Unk
Blue Canyon TechnologiesRW84.400100.2508.0001Unk
CubeSpace Satellite SystemsCubeWheel Small0.0600.650.000230.00177124
CubeSpace Satellite SystemsCubeWheel Small+0.0902.30.00230.0036124
CubeSpace Satellite SystemsCubeWheel Medium0.1502.30.0010.01082124
CubeSpace Satellite SystemsCubeWheel Large0.2254.50.00230.03061124
GomSpaceNanoTorque GSW-6000.9400.30.00150.0191Unk
ComatRW200.18010.0020.021Up to 20Krad*
ComatRW400.23010.0040.041Up to 20Krad*
ComatRW600.27510.0060.061Up to 20Krad*
AAC Clyde SpaceRW2100.480.80.00010.006136
AAC Clyde SpaceRW4000.375150.0080.050136
AAC Clyde SpaceTrillian-11.52447.11.21Unk
NanoAvionicsRWO0.1373.250.00320.020120
NanoAvionics4RWO0.66560.00590.037420
NewSpace SystemsNRWA-T6<51360.30.00783120
NewSpace SystemsNRWA-T0651.551.70.020.00094110
NewSpace SystemsNRWA-T22.80.40.090.00163110
Rocket LabRW-0.030.1851.80.0020.040120
Rocket LabRW-0.0030.048Unk0.0010.005110
Rocket LabRW-0.010.1221.050.0010.018120
Rocket LabRW3-0.060.23523.40.0200.180120
Rocket LabRW4-0.20.6Unk0.10.2160
Rocket LabRW4-0.40.77Unk0.10.4160
Rocket LabRW4-1.01.38430.11160
Vectronic AerospaceVRW-A-11.901100.0906.000120
Vectronic AerospaceVRW-B-21.00450.0200.200120
Vectronic AerospaceVRW-C-12.3450.0201.20120
Vectronic AerospaceVRW-D-22650.052.0120
Vectronic AerospaceVRW-D-631100.096120
*Printed Circuit Board (PCB) level

5.2.3 Magnetic Torquers

Magnetic torquers provide control torques perpendicular to the local external magnetic field. Table 5-4 lists a selection of high heritage magnetic torquers and figure 5.3 illustrates some of ZARM Technik’s product offerings. Magnetic torquers are often used to remove excess momentum from reaction wheels. As control torques can only be provided in the plane perpendicular to the local magnetic field, magnetic torquers alone cannot provide three-axis stabilization.

Use of magnetic torquers beyond low-Earth orbit and in interplanetary applications need to be carefully investigated since their successful operation is relying on a significant local external magnetic field. This magnetic field may or may not be available in the location and environment for that mission and additional control methods may be required.

photo of hardware
Figure 5.3: Magnetorquers for micro satellites.
Credit: ZARM Technik
Table 5-4: High Heritage Magnetic Torquers
ManufacturerModelMass (kg)Power (W)Peak Dipole
(A m2)
# AxesRadiation Tolerance (krad)
CubeSpace Satellite SystemsCubeTorquer Small0.0280.420.24124
CubeSpace Satellite SystemsCubeTorquer Medium0.0360.370.66124
CubeSpace Satellite SystemsCubeTorquer Large0.0720.371.90124
CubeSpace Satellite SystemseCubeTorquer Coil(Single)0.0460.310.13124
CubeSpace Satellite SystemsCubeTorquer Coil(Double)0.0740.640.27124
GomSpaceNano Torque GST-6000.156Unk0.31 – 0.343Unk
GomSpaceNanoTorque Z-axis Internal0.106Unk0.1391Unk
ISISPACEMagnetorquer Board0.1961.20.203Unk
MEISEIMagnetic Torque Actuator for Spacecraft0.51121Unk
AAC Clyde SpaceMTQ8000.3953151Unk
NanoAvionicsMTQ3X0.2050.40.30320
NewSpace SystemsNCTR-M0030.0300.250.291Unk
NewSpace SystemsNCTR-M0120.0530.81.191Unk
NewSpace SystemsNCTR-M0160.0531.21.61Unk
Rocket LabTQ-400.825Unk48.001Unk
Rocket LabTQ-150.400Unk19.001Unk
ZARM Technik**MT0.2-10.012-0.0140.135-0.250.21NA*
ZARM TechnikMT0.5-10.0090.2750.51NA*
ZARM TechnikMT0.7-1-010.0350.50.71NA*
ZARM TechnikMT1-1-010.0650.2311NA*
ZARM TechnikMT1.5-1-010.0970.41.51NA*
ZARM TechnikMT2-1-020.10.521NA*
ZARM TechnikMT3-1-D220427010.150.731NA*
ZARM TechnikMT4-10.150.641NA*
ZARM TechnikMT5-10.19-0.30.73-0.7551NA*
ZARM TechnikMT5-20.310.7751NA*
ZARM TechnikMT6-20.25-0.30.48-1.161NA*
ZARM TechnikMT7-20.40.971NA*
ZARM TechnikMT10-10.35-0.40.53-0.8101NA*
ZARM TechnikMT10-20.37-0.480.7-1101NA*
ZARM TechnikMT15-10.4-0.551.0-1.55151NA*
ZARM TechnikMT15-20.5-0.550.9-1.5151NA*
* Only EEE parts are connector and wires. Magnetorquer is not sensitive to ionizing radiation.
** ZARM Technik: Over 200 models available with design to mass/power optimization

5.2.4 Thrusters

Thrusters used for attitude control are described in Chapter 4: In-Space Propulsion. Pointing accuracy is determined by minimum impulse bit, and control authority by thruster force.

5.2.5 Star Trackers

A star tracker can provide an accurate estimate of the absolute three-axis attitude by comparing a digital image to an onboard star catalog (8). Star trackers identify and track multiple stars and provide three-axis attitude several times a second. Table 5-5 lists some models suitable for use on small spacecraft. For example, Arcsec’s Sagitta Star Tracker was launched on the SIMBA cubesat in 2020.

Table 5-5: Star Trackers Suitable for Small Spacecraft
ManufacturerModelMass (kg)Power (W)FOVCross axis accuracy (3s)Twist accuracy
(3s)
Radiation Tolerance (krad)TRL
Redwire SpaceStar Tracker0.4752.514×1910/27″51″757-9
ArcsecSagitta0.2751.425.4°630207-9
ArcsecTwinkle0.040.610.4°30180Unk7-9
Ball AerospaceCT-20203.0008Unk1.5”1”Unk5-6
Berlin Space Technologies / AAC Clyde SpaceST2000.0400.6522°30″200″117-9
Berlin Space Technologies / AAC Clyde SpaceST4000.2500.7515°15″150″117-9
Blue Canyon TechnologiesStandard NST0.3501.510° x 12°6″40″Unk7-9
Blue Canyon TechnologiesExtended NST1.3001.510° x 12°6″40″Unk7-9
CreareUST0.840UnkUnk7″15″Unk5-6
CubeSpace Satellite SystemsCubeStar0.0550.26458-47°55.44″ 0.02°77.4197-9
Danish Technical UniversityMicroASC0.4251.9Unk2”UnkUnk7-9
LeonardoSpacestar1.600620° x 20°7.7″10.6″Unk7-9
NanoAvionicsST-10.1081.221° full-cone8″50″207-9
Rocket LabST-16RT20.18518° half-cone5″55″Unk7-9
SodernAuriga-CP0.2051.1Unk2″11″Unk7-9
SodernHydra-M2.757UnkUnkUnkUnk5-6
SodernHydra-TC5.38UnkUnkUnkUnk5-6
Solar MEMS TechnologiesSTNS0.14112°40″70″207-9
Space MicroMIST0.520314.5°15″105″307-9
Space MicroµSTAR-100M1.8005Unk15″105″100Unk
Space MicroµSTAR-200M2.1008-10Unk15″105″100Unk
Space MicroµSTAR-200H2.70010Unk3″21″100Unk
Space MicroµSTAR-400M3.30018Unk15″105″100Unk
TermaT10.760.820° circular2.2″9″1005-6
TermaT30.35220° circular2.6″10″85-6
Vectronic AerospaceVST-41MN0.7 – 0.92.514° x 14°27″183″207-9
Vectronic AerospaceVST-68M0.470314° x 14°7.5″45″20Unk

5.2.6 Magnetometers

Magnetometers provide a measurement of the local magnetic field which can be used to estimate 2-axis information about the attitude (9). Table 5-6 provides a summary of some three-axis magnetometers available for small spacecraft, one of which is illustrated in figure 5.4.

photo of hardware
Figure 5.4: NSS Magnetometer.
Credit: NewSpace Systems
Table 5-6: Three-axis Magnetometers for Small Spacecraft
ManufacturerModelMass (kg)Power (W)Resolution (nT)OrthogonalityRadiation Tolerance (krad)TRL
GomSpaceNanoSense M3150.008UnkUnkUnkUnk7-9
AAC Clyde SpaceMM2000.0120.011.18Unk307-9
MEISEI3-Axis Magnetometer for Small Satellite0.2201.5UnkUnk7-9
NewSpace SystemsNMRM-Bn25o4850.0850.758107-9
AAC Clyde SpaceMAG-30.100Voltage DependentUnk107-9
ZARM TechnikAnalogue High-Rel Fluxgate Magnetometer FGM-A-750.330.75 W±75000509
ZARM TechnikDigital AMR Magnetometer AMR-D-100-EFRS4850.180.3 W±100000unk6-7

5.2.7 Sun Sensors

Sun sensors are used to estimate the direction of the Sun in a spacecraft body frame. Sun direction estimates can be used for attitude estimation, though to obtain a three-axis attitude estimate at least one additional independent source of attitude information is required (e.g., the Earth nadir vector or the direction to a star). Because the Sun is easily identifiable and extremely bright, Sun sensors are often used for fault detection and recovery. However, care must be taken to ensure the Moon or Earth’s albedo is not inadvertently perturbing the measurement.

photo of hardware
Figure 5.5: Redwire Coarse Sun Sensor Detector (Cosine Type).
Credit: Redwire Space

There are several types of Sun sensors which operate on different principles.

Cosine detectors are photocells. Their output is the current generated by the cell, which is (roughly) proportional to the cosine of the angle between the sensor boresight and the Sun. Typically several cosine detectors (pointing in different directions) are used on a spacecraft for full sky coverage. Cosine detectors (e.g., figure 5.5) are inexpensive, low-mass, simple and reliable devices, but their accuracy is typically limited to a few degrees, and they do require analog-to-digital converters.

Quadrant detectors. Quadrant sun sensors typically operate by shining sunlight through a square window onto a 2 x 2 array of photodiodes. The current generated by each photodiode is a function of the direction of the Sun relative to the sensor boresight. The measured currents from all four cells are then combined mathematically to produce the angles to the Sun.

Digital Sun Sensor. The Sun illuminates a narrow slit behind which, is located a geometric coded bit mask and a number of photodiodes under the mask. Depending on the angle to the Sun, the photodiodes will be illuminated as per the geometric pattern resulting in correspondingly different photocurrents which are then amplified and thresholded against an average value. Given the known slit geometries, this digital bit output can be then converted to a sun angle.

Sun Camera. Some sun sensors are build as a small camera imaging the Sun. Since the Sun is so bright, the optics will include elements to decrease the throughput. A computer will identify the image of the Sun and calculate the centroid. Sun sensors can be made very accurate this way. Sometimes, multiple apertures are included to increase accuracy.

Examples of small spacecraft sun sensors are described in table 5-7.

Table 5-7: Small Spacecraft Sun Sensors
ManufacturerModelSensor TypeMass (kg)Peak Power (W)Analog or DigitalFOVAccuracy (3s)# Measurement AnglesRadiation Tolerance (krad)TRL
Redwire SpaceCoarse Analog Sun SensorCoarse Analog Sun Sensor0.0450Analog±40° (Can be modified to meet specific FOV requirements)±1°1>1007-9
Redwire SpaceCoarse Sun Sensor (Cosine Type)Coarse Sun Sensor (Cosine Type)0.0100AnalogAPPROXIMATE COSINE, CONICAL SYMMETRY±2° to ±5°Depends on configuration>1007-9
Redwire SpaceCoarse Sun Sensor PyramidCoarse Sun Sensor Pyramid0.13
0
Analog2π STERADIAN PLUS ±1° to ±3°2>1007-9
Redwire SpaceDIGITAL SUN SENSOR (±32°)DIGITAL SUN SENSOR (±32°)Sensor
0.3 kg
Electronics
~1
1Digital±32° x ±32° (each sensor)±0.125°21007-9
Redwire SpaceDigital Sun Sensor (±64°)Digital Sun Sensor (±64°)Sensor0.25
Electronics 0.29 – 1.1
0.5Digital128° X 128° (EACH SENSOR)
NOTE: 4π STERADIANS
ACHIEVED WITH 5 SENSORS
±0.25°21007-9
Redwire SpaceFine Pointing Sun SensorFine Pointing Sun SensorSensor .95
Electronics 1.08
< 3Digital±4.25° x ±4.25° (Typical)Better than ±0.01°21007-9
Redwire SpaceFine Spinning Sun Sensor (±64°)Fine Spinning Sun Sensor (±64°)Sensor 0.109
Electronics
0.475 – 0.725
0.5Analog and Digital±64° FAN SHAPED (each sensor)±0.1°1
plus Sun Pulse
1007-9
Redwire SpaceMicro Sun SensorMicro Sun Sensor< 0.002 < 0.02Analog± 85° MINIMUM±5°2Approx. 105-6
Redwire SpaceMiniature Spinning Sun Sensor (±87.5°)Miniature Spinning Sun Sensor (±87.5°)< 0.250.5Digital±87.5° (FROM NORMAL TO SPIN AXIS)±0.1°1
plus Sun Pulse
1007-9
Redwire SpaceFINE SUN SENSOR (±50°)FINE SUN SENSOR (±50°)UnkUnkDigital100 X 100 Each Sensor±0.01° TO ±0.05°2100, 150, or 3007-9
Bradford SpaceCoSSCosine0.0240Analog160° full cone1400007-9
Bradford SpaceCoSS-RCosine0.0150Analog180° full cone11200007-9
Bradford SpaceCSS-01, CSS-02 Only shows one CSSCosine0.2150Analog180° full cone1.5°2700007-9
Bradford SpaceFSSQuadrant0.3750.25Analog128° x 128°0.3°21007-9
Bradford SpaceMini-FSSQuadrant0.0500Analog128° x 128°0.2° With on-board implementation2200007-9
CubeSpace Satellite SystemsCubeSenseCamera0.030<0.2Digital180°0.2°2247-9
GomSpaceNanoSense FSSQuadrant0.002UnkDigital{45°, 60°}{±0.5°, ±2°}2UnkUnk
AAC Clyde SpaceSS200Unk.0030.04Digital110°<1°Unk>367-9
Lens R&DBiSon64-ETQuadrant0.0230Analog±58° per axis0.5°292007-9
Lens R&DBiSon64-ET-BQuadrant0.0330Analog±58° per axis0.5°292007-9
Lens R&DMAUSQuadrant0.0140Analog±57° per axis0.5°292007-9
NewSpace SystemsNFSS-411Unk0.0350.150Digital140°0.1°TBD207-9
NewSpace SystemsNCSS-SA05Unk0.0050.05Analog114°0.5°TBDUnk7-9
Solar MEMS TechnologiesnanoSSOC-A60Orthogonal0.0040.007Analog±60° per axis0.5°21007-9
Solar MEMS TechnologiesnanoSSOC-D60Orthogonal0.0070.076Digital±60° per axis0.5°2307-9
Solar MEMS TechnologiesSSOC-A60Orthogonal0.0250.01Analog±60° per axis0.5°21007-9
Solar MEMS TechnologiesSSOC-D60Orthogonal0.0350.315Digital±60° per axis0.5°2307-9
Solar MEMS TechnologiesACSSQuadrant & Redundant0.0350.072Analog±60° per axis0.5°22007-9
Space MicroCSS-01, CSS-02Cosine0.0100Analog120° full cone11007-9
Space MicroMSS-01Quadrant0.0360Analog48° full cone21007-9

5.2.8 Horizon Sensors

Horizon sensors can be simple infrared horizon crossing indicators (HCI), or more advanced thermopile sensors that can detect temperature differences between the poles and equator. For terrestrial applications, these sensors are referred to as Earth Sensors, but can be used for other planets. Examples of such technologies are described in table 5-8 and illustrated in figure 5.6.

In addition to the commercially-available sensors listed in table 5-8, there has been some recent academic interest in horizon sensors for CubeSats with promising results (24) (10) (11).

photo of hardware
Figure 5.6: MAI-SES.
Credit: Redwire Space
Table 5-8: Commercially Available Horizon Sensors
ManufacturerModelSensor TypeMass (kg)Peak Power (W)Analog or DigitalAccuracy# Measurement AnglesRad Tolerance (krad)TRL
CubeSpace Satellite SystemsCubeSenseCamera0.0300.200Digital0.2°2247-9
ServoMini Digital HCIPyroelectric0.050Voltage DependentDigital0.75°UnkUnk7-9
ServoRH 310 HCIPyroelectric1.51Unk0.015°Unk20Unk
SITAELDigital Earth SensorMicrobolometer0.4<2Digital<1°UnkUnkUnk
Solar MEMS TechnologiesHSNSInfrared0.1200.150Digital2307-9

5.2.9 Inertial Sensing

Inertial sensors include gyroscopes for measuring angular change and accelerometers for measuring velocity change. They are packaged in different ways that range from single-axis devices (i.e., a single gyroscope or accelerometer), to packages which include 3 orthogonal axes of gyroscopes (Inertial Reference Unit (IRU)) to units containing 3 orthogonal gyros and 3 orthogonal accelerometers (Inertial Measurement Unit (IMU)). These sensors are frequently used to propagate the vehicle state between measurement updates of a non-inertial sensor. For example, star trackers typically provide attitude updates at a few Hertz. If the control system requires accurate knowledge between star tracker updates, then an IMU may be used for attitude propagation between star tracker updates.

Gyroscope technologies typically used in modern small spacecraft are fiber optic gyros (FOGs) and MEMS gyros, with FOGs usually offering superior performance at a mass and cost penalty (12). Other gyroscope types exist (e.g., resonator gyros, ring laser gyros), but these are not common in the SmallSat/CubeSat world due to size, weight, and power (SWaP) and cost considerations.

Gyro behavior is a complex topic (13) and gyro performance is typically characterized by a multitude of parameters. Table 5-9 only includes bias stability and angle random walk for gyros, and bias stability and velocity random walk for accelerometers, as these are often the driving performance parameters. That said, when selecting inertial sensors, it is important to consider other factors such as dynamic range, output resolution, bias, sample rate, etc.

Table 5-9: Gyros Available for Small Spacecraft
ManufacturerModelSensor TypeTechnologyMass (kg) Power (W)GyrosAccelerometers
 Bias StabilityARW Bias StabilityVRW
# Axes(°/hr)stat (°/rt(hr))# Axes (µg)stat(m/sec)/rt(hr)
EmcoreQRS11GyroMEMS≤0.060.816TypicalN/AN/AN/AN/AN/A
EmcoreQRS28GyroMEMS≤0.0250.52N/AN/AN/AN/AN/AN/AN/A
HoneywellMIMUIMURLG43430.05Unk0.01Unk100UnkUnk
HoneywellHG1700IMURLG0.95.00031.0001s0.125310001s0.65
L3CIRUSGyrosFOG15.40040.00030.0001s0.1000N/AUnkN/A
NewSpace SystemsNSGY-001IRUImage-based rotation estimate0.0550.2003N/AN/A0N/AUnkN/A
Northrop GrummanLN-200SIMUFOG, SiAc0.7481231.0001s0.07033001sUnk
NovAtelOEM-IMU-STIM300IMUMEMS0.0551.5030.500TBD0.150350TBD0.060
SafranSTIM202IRUMEMS0.0551.50030.400TBD0.1700N/ATBDN/A
SafranSTIM210IRUMEMS0.0521.50030.300TBD0.1500N/ATBDN/A
SafranSTIM300IMUMEMS0.0552.00030.300TBD0.150350TBD0.07
SafranSTIM318IMUMEMS0.0572.50030.300TBD0.15033TBD0.015
SafranSTIM320IMUMEMS0.0572.50030.300TBD0.10033TBD0.015
SafranSTIM277HIRUMEMS0.0521.50030.300TBD0.1500N/ATBDN/A
SafranSTIM377HIMUMEMS0.0552.00030.300TBD0.150350TBD0.07
Silicon Sensing SystemsCRH03GyroMEMS0.420.2W1CRH03-010 –0.03
CRH03-025 – 0.04
CRH03-100 – 0.04
CRH03-200 – 0.05
CRH03-400 – 0.1
 CRH03-010 – 0.005
CRH03-025 – 0.006
CRH03-100 – 0.006
CRH03-200 – 0.008
CRH03-400 – 0.010
0N/AN/A
Silicon Sensing SystemsCRH03
(OEM)
GyroMEMS0.180.2W1CRH03-010 – 0.03
CRH03-025 – 0.04
CRH03-100 – 0.04
CRH03-200 – 0.05
CRH03-400 – 0.1
 CRH03-010 – 0.005
CRH03-025 – 0.006
CRH03-100 – 0.006
CRH03-200 – 0.008
CRH03-400 – 0.010
0N/AN/A
Silicon Sensing SystemsRPU30GyroMEMS1.35<0.8W30.06 0.0060N/AN/A
Silicon Sensing SystemsDMU419 DoF IMUMEMS<2<1.5W30.1 0.0153150.05
Silicon Sensing SystemsCASAccMEMS0.004Unk0N/A N/A2CAS2X1S – 7.5
CAS2X2S – 7.5
CAS2X3S – 7.5
CAS2X4S – 25
CAS2X5S – 75
 CAS2X1S – TBC
CAS2X2S – TBC
CAS2X3S – TBC
CAS2X4S – TBC
CAS2X5S – TBC
VectorNavVN-100*IMU + magnetometers +barometerMEMS0.0150.220310.000max0.210340max0.082
VectorNavVN-110*IMU + magnetometersMEMS0.1252.50031.000max0.0833310max0.024
*Small form-factor versions of these products available.

5.2.10 GPS Receivers

For low-Earth orbit spacecraft, GPS receivers are now the primary method for performing orbit determination, replacing ground-based tracking methods. Onboard GPS receivers are now considered a mature technology for small spacecraft, and some examples are described in table 5-10. There are also next-generation chip-size COTS GPS solutions, for example the NovaTel OEM 719 board has replaced the ubiquitous OEMV1.

GPS accuracy is limited by propagation variance through the exosphere and the underlying precision of the civilian use C/A code (14). GPS units are controlled under the Export Administration Regulations (EAR) and must be licensed to remove Coordinating Committee for Multilateral Export Control (COCOM) limits (15).

Although the usability of GPS is limited to LEO missions, past experiments have demonstrated the ability of using a weak GPS signal at GSO, and potentially soon to cislunar distances (16) (17). Development and testing in this fast-growing area of research and development may soon make onboard GPS receivers more commonly available.

Table 5-10: GPS Receivers for Small Spacecraft
ManufacturerMass (kg)Power (W)Accuracy (m)Radiation Tolerance (krad)TRL
AAC Clyde Space0.16Unk<5107-9
APL0.41.415205-6
General Dynamics1.2815Unk7-9
General Dynamics1.1815Unk7-9
GomSpace0.0311.31.5187-9
SkyFox Labs0.0240.12410307-9
Spacemanic0.025~0.11.5407-9
Surrey Satellite Technology0.090.5557-9
Syrlinks0.435Unk<0.1155-6

5.2.11 Deep Space Navigation

In deep space, navigation is performed using radio transponders in conjunction with the Deep Space Network (DSN). As of 2020, the only deep space transponder with flight heritage suitable for small spacecraft was the JPL-designed and General Dynamics-manufactured Small Deep Space Transponder (SDST). JPL has also designed IRIS V2, which is a deep space transponder that is more suitable for the CubeSat form factor. Table 5-11 details these two radios, and the SDST is illustrated in figure 5.7. IRIS V2, derived from the Low Mass Radio Science Transponder (LMRST), has flown on the MarCO CubeSats in 2018, LICIACube that performed an asteroid flyby in September 2022, 12U lunar CAPSTONE spacecraft that entered a lunar orbit November 13, 2022, and was on six Artemis 1 secondary CubeSat payloads (Lunar Flashlight, LunaH-Map, ArgoMoon, CubeSat for Solar Particles, Biosentinel, and NEA Scout). It is also scheduled to fly on INSPIRE (18).

photo of hardware
Figure 5.7: General Dynamics SDST.
Credit: General Dynamics
Table 5-11: Deep Space Transponders for Small Spacecraft
ManufacturerModelMass (kg)Rx Power (W)BandsRadiation Tolerance (krad)TRL
General DynamicsSDST3.212.5X, Ka507-9
Space Dynamics LaboratoryIRIS V2.11.110.3X, Ka, S, UHF157-9

5.2.12 Atomic Clocks

Atomic clocks have been used on larger spacecraft in low-Earth orbit for several years now, however integrating them on small spacecraft is relatively new. Table 5-12 provides examples of commercially available atomic clocks and oscillators for SmallSats. The conventional method for spacecraft navigation is a two-way tracking system of ground-based antennas and atomic clocks. The time difference from a ground station sending a signal and the spacecraft receiving the response can be used to determine the spacecraft’s location, velocity, and (using multiple signals) the flight path. This is not a very efficient process, as the spacecraft must wait for navigation commands from the ground station instead of making real-time decisions, and the ground station can only track one spacecraft at a time, as it must wait for the spacecraft to return a signal (19). In deep space navigation, the distances are much greater from the ground station to spacecraft, and the accuracy of the radio signals needs to be measured within a few nanoseconds.

More small spacecraft designers are developing their own version of atomic clocks and oscillators that are stable and properly synchronized for use in space. They are designed to fit small spacecraft, for missions that are power- and volume-limited or require multiple radios.

Table 5-12: Atomic Clocks and Oscillators for Small Spacecraft
ManufacturerModelDimensions (mm)Mass (kg)Power (W)Frequency Range (MHz)Rad Tolerance (krad)TRL
AccuBeatUltra Stable Oscillator131 x 120 x 10526.557.51852507-9
Bliley Technologies Iris Series 1″x1″ OCXO for LEO19 x 11 x 190.0161.510 -100397-9
Aether Series TCVCXO for LEO21 x 14 x 8Unk0.05610 – 15037Unk
MicrosemiSpace Chip Scale Atomic Clock (CSAC)41 x 36 x 120.0350.1210205-6
Safran Timing Technologies SAMO44 x 54 x 570.223.5 Nom 5.5 Max10100Unk
Space Qualified mRO-5051 x 51 x 200.0800.4 Nom1025 MinUnk
miniRAFS108 x 53 x 680.45< 12 Max60 and 10TBDUnk
LNMO50 x 50 x 300.11.5 Nom 2.5 Max5 – 40100Unk

5.2.13 LiDAR

Light Detection and Ranging (LiDAR) is new type of sensor that is emerging. The technology has matured in terrestrial applications (such as automotive applications) over the last decade and is used in larger spacecraft that are capable of proximity operations, like Orion. This sensor type has applications for small spacecraft altimetry and relative navigation (e.g., a Mars helicopter, rendezvous and docking, and formation flying). Table 5-13 lists examples of flown LiDARs.

Table 5-13: Lidar for Small Spacecraft
ManufacturerModelMass (kg)Power (W)Max Range (m)Radiation Tolerance (krad)TRL
GarminLidar Lite V30.0220.740Unk5‑6*
ASCGSFL-4K (3D)330>1 km in altimeter modeUnk7-9
*Specific units were qualified for Mars Ingenuity helicopter. Product line in general is not space qualified.

5.3 On the Horizon

In general, technological progress in guidance, navigation, and control is advancing quickly in automotive research areas but is lagging slightly in the aerospace industry. Given the high maturity of existing GNC components, future developments in GNC are mostly focused on incremental or evolutionary improvements, such as decreases in mass and power, and increases in longevity and/or accuracy. This is especially true for GNC components designed for deep space missions that have only very recently been considered for small spacecraft. However, in a collaborative effort between the Swiss Federal Institute of Technology and Celeroton, there is progress being made on a high-speed magnetically levitated reaction wheel for small satellites (figure 5.8). The idea is to eliminate mechanical wear and stiction by using magnetic bearings rather than ball bearings. The reaction wheel implements a dual hetero/homopolar, slotless, self-bearing, permanent-magnet synchronous motor (PMSM). The fully active, Lorentz-type magnetic bearing consists of a heteropolar self-bearing motor that applies motor torque and radial forces on one side of the rotor’s axis, and a homopolar machine that exerts axial and radial forces to allow active control of all six degrees of freedom. It can store 0.01 Nm of momentum at a maximum of 30,000 rpm, applying a maximum torque of 0.01 Nm (21).

Several projects funded via NASA’s Small Spacecraft Technology (SST) program through the University Smallsat Technology Partnerships (USTP) initiative have begun advancing GNC systems. Listed below in table 5-14 are projects that focused on GNC advancement, and further information can be found at the USTP website:

https://www.nasa.gov/smallspacecraft/university-smallsat-technology-partnership-initiative/

Each presentation is from the USTP Technology Exposition that was held in May 2021 and June 2022.

photo of hardware
Figure 5.8: High-speed magnetically levitated reaction wheel.
CreditL Celeroton AG
Table 5-14: USTP Initiative GNC Projects
ProjectUniversityCurrent StatusReference
On-Orbit Demonstration of Surface Feature-Based Navigation and TimingUniversity of Texas, AustinStill in developmentUSTP Technology Expo presentation
Autonomous Nanosatellite Swarming (ANS) using Radio Frequency and Optical NavigationStanford UniversityOnboard Starling mission (Launched in 2023)USTP Technology Expo presentation
Distributed multi-GNSS Timing and Localization (DiGiTaL)Stanford UniversityLeveraged technology used in Starling missionUSTP Technology Expo presentation
Mems Reaction Control and Maneuvering for Picosat beyond LEOPurdue UniversityAwarded a suborbital flight test through NASA’s Flight Opportunities program(29)
A Small Satellite Lunar Communications and Navigation SystemUniversity of Boulder, ColoradoStill in developmentUSTP Technology Expo presentation
A high-precision continuous-time PNT compact module for the LunaNet small spacecraftUniversity of California, Los AngelesStill in developmentUSTP Technology Expo presentation

5.4 Summary

Conventional small spacecraft GNC technology is a mature area, with many high TRL components previously flown around Earth offered by several different vendors. These GNC techniques are generally semi/non-autonomous as on-board observations are collected with the assistance of ground-based intervention. As the interest for deep space exploration with small spacecraft grows, semi-to-fully autonomous navigation methods must advance. It is likely that future deep space navigation will rely solely on fully autonomous GNC methods that require zero ground-based intervention to collect/provide navigation data. This is a desirable capability as the spacecraft’s dependence on Earth-based tracking resources (such as DSN) is reduced and the demand for navigation accuracy increases at large distances from Earth. However, current methods advancing deep space navigation involve both ground- and space-based tracking in conjunction with optical navigation techniques. To support this maturity, the small spacecraft industry has seen a spike in position, navigation, and timing (PNT) technology progression in inertial sensors and atomic clocks, and magnetic navigation for near-Earth environments.

Other GNC advances involve research on SmallSats performing on-orbit proximity operations. Several research papers have discussed ways to accomplish this, and previous extravehicular free flyers have demonstrated this innovative capability in the past few decades. The CubeSat Proximity Operations Demonstration (CPOD) project is the most recent CubeSat mission to attempt the characterization of low-power proximity operations technologies, however its mission ended June 2023 and was unable to demonstrate rendezvous, proximity operations and docking maneuvers as planned. Seeker, a 3U CubeSat that was deployed September 2019, was built to demonstrate safe operations around a target spacecraft with core inspection capabilities. While Seeker was unable to perform its underlining goal, there were still several benefits for improving future missions (29).

The rising popularity of SmallSats in general, and CubeSats in particular, means there is a high demand for components, and engineers are often faced with prohibitive prices. The Space Systems Design Studio at Cornell University is tackling this issue for GNC with their PAN nanosatellites. A paper by Choueiri et al. outlines an inexpensive and easy-to-assemble solution for keeping the ADCS system below $2,500 (22). Lowering the cost of components holds exciting implications for the future and will likely lead to a burgeoning of the SmallSat industry.

For feedback solicitation, please email: arc-sst-soa@mail.nasa.gov. Please include a business email so someone may contact you further.

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