In this paper, first ever flight demonstrations of Optimal Propellant Maneuver (OPM), a method of propulsive rotational state transition for spacecraft controlled using thrusters, is presented for the International Space Station (ISS). On August 1, 2012, two ISS reorientations of about 180deg each were performed using OPMs. These maneuvers were in preparation for the same-day launch and rendezvous of a Progress vehicle, also a first for ISS visiting vehicles. The first maneuver used 9.7 kg of propellant, whereas the second used 10.2 kg. Identical maneuvers performed without using OPMs would have used approximately 151.1kg and 150.9kg respectively. The OPM method is to use a pre-planned attitude command trajectory to accomplish a rotational state transition. The trajectory is designed to take advantage of the complete nonlinear system dynamics. The trajectory choice directly influences the cost of the maneuver, in this case, propellant. For example, while an eigenaxis maneuver is kinematically the shortest path between two orientations, following that path requires overcoming the nonlinear system dynamics, thereby increasing the cost of the maneuver. The eigenaxis path is used for ISS maneuvers using thrusters. By considering a longer angular path, the path dependence of the system dynamics can be exploited to reduce the cost. The benefits of OPM for the ISS include not only reduced lifetime propellant use, but also reduced loads, erosion, and contamination from thrusters due to fewer firings. Another advantage of the OPM is that it does not require ISS flight software modifications since it is a set of commands tailored to the specific attitude control architecture. The OPM takes advantage of the existing ISS control system architecture for propulsive rotation called USTO control mode1. USTO was originally developed to provide ISS Orbiter stack attitude control capability for a contingency tile-repair scenario, where the Orbiter is maneuvered using its robotic manipulator relative to the ISS. Since 2005 USTO has been used for nominal ISS operations

Bedrossian, Nazareth

Bhatt, Sagar

Longacre, Kenneth

Nguyen, Louis

NASA Technical Reports Server

 
American Institute of Aeronautics and Astronautics 
 
 
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Optimal Propellant Maneuver Flight Demonstrations on ISS 
Sagar Bhatt1 
The Charles Stark Draper Laboratory, Inc., Houston, TX 77058 
Nazareth Bedrossian2 
Halliburton, Houston, TX 
Kenneth Longacre3 
United Space Alliance,LLC, Houston, TX 77058 
Louis Nguyen4 
NASA Johnson Space Center, Houston, TX 77058 
This paper presents results for the flight demonstrations of Optimal Propellant 
Maneuver (OPM) on August 1, 2012. The OPM was used to complete two docking 
maneuvers of the International Space Station (ISS) using only 19.9 kg of propellant, saving 
93% compared to using ISS flight software. The savings were achieved by commanding the 
ISS to follow a pre-planned attitude trajectory which was optimized to take advantage of 
naturally occurring environmental torques and available control authority from the jets. 
The trajectory was obtained by solving an optimal control problem. The flight 
implementation did not require any modifications to flight software. This approach is 
applicable to any spacecraft controlled with thrusters. 
I. Introduction 
N this paper, first ever flight demonstrations of Optimal Propellant Maneuver (OPM), a method of propulsive 
rotational state transition for spacecraft controlled using thrusters, is presented for the International Space Station 
(ISS). On August 1, 2012, two ISS reorientations of about 180deg each were performed using OPMs. These 
maneuvers were in preparation for the same-day launch and rendezvous of a Progress vehicle, also a first for ISS 
visiting vehicles. The first maneuver used 9.7 kg of propellant, whereas the second used 10.2 kg. Identical 
maneuvers performed without using OPMs would have used approximately 151.1kg and 150.9kg respectively. 
 The OPM method is to use a pre-planned attitude command trajectory to accomplish a rotational state transition. 
The trajectory is designed to take advantage of the complete nonlinear system dynamics. The trajectory choice 
directly influences the cost of the maneuver, in this case, propellant. For example, while an eigenaxis maneuver is 
kinematically the shortest path between two orientations, following that path requires overcoming the nonlinear 
system dynamics, thereby increasing the cost of the maneuver. The eigenaxis path is used for ISS maneuvers using 
thrusters. By considering a longer angular path, the path dependence of the system dynamics can be exploited to 
reduce the cost. 
The benefits of OPM for the ISS include not only reduced lifetime propellant use, but also reduced loads, 
erosion, and contamination from thrusters due to fewer firings. Another advantage of the OPM is that it does not 
require ISS flight software modifications since it is a set of commands tailored to the specific attitude control 
architecture. 
The OPM takes advantage of the existing ISS control system architecture for propulsive rotation called USTO 
control mode1. USTO was originally developed to provide ISS Orbiter stack attitude control capability for a 
contingency tile-repair scenario, where the Orbiter is maneuvered using its robotic manipulator relative to the ISS. 
Since 2005 USTO has been used for nominal ISS operations. 
                                                          
1 Member of Technical Staff II, Vehicle Dynamics & Control Group. 
2 Chief Technical Advisor – Mechanical, Wireline Systems. 
3 Attitude Determination and Control Officer, Mission Operations Directorate. 
4 Deputy Manager, Space Station GN&C, Engineering Directorate. 
I
https://ntrs.nasa.gov/search.jsp?R=20130014033 2019-08-29T15:30:40+00:00Z
 
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Previous work on minimum propellant spacecraft rotations has utilized bang-off-bang profiles2 designed based 
on simplifying assumptions such as rest-to-rest maneuvers, inertially symmetric spacecraft, and no disturbance 
torques3. A three-pulse algorithm with reduced propellant usage has also been developed for axisymmetric 
spacecraft with non-zero initial rates4. Recent computational advances have allowed the solution of increasingly 
complex optimization problems. Consequently, the minimum fuel problem can be solved with fewer simplifying 
assumptions. By incorporating the full nonlinear dynamics, disturbance torques, and a varying control profile, the 
OPM trajectory can be shaped to complete the rotation as efficiently as possible. 
OPM is similar to the Zero-Propellant Maneuver (ZPM), which commanded the CMGs to complete the 
maneuver while simultaneously performing momentum dumping. The first ZPM flight demonstration on November 
5, 2006 rotated the ISS 90 degrees5. On March 3, 2007, a second ZPM demonstration rotated the ISS 180 degrees 
without using propellant6. To avoid CMG saturation, ZPM generally requires longer maneuver times. With OPM, 
the maneuver times can be made significantly shorter, depending on the propellant budget, thus allowing fewer 
restrictions on ISS operations and requiring less ISS thermal analysis. Another advantage of OPM compared to ZPM 
is that the ISS thrusters have much more control authority than the CMGs and thus can better handle discrepancies 
between predicted and on-orbit initial conditions, mass properties, and environment. 
 This paper presents the OPM method within the context of two specific ISS docking maneuvers. The ISS 
environmental dynamics, control systems, and operational modes are introduced, and a suitable optimal control 
problem solved. Specific constraints imposed on the control problem due to ISS flight software and operations will 
be presented. The flight operational maneuver is described and OPM trajectory is verified in high fidelity 
simulation. Flight results are then presented. 
II. Problem Formulation and Solution Technique 
A rotational state transition, e.g. attitude, rate and angular momentum, can be planned to minimize a user 
specified “cost” by posing and solving an optimal control problem (OCP) for a specified maneuver time. The OCP 
in this context is to transition the spacecraft from an initial to a final rotational state satisfying the system dynamics 
and any constraints on the system.  For the ISS, the system dynamics include Euler, gravity gradient and 
aerodynamic torques, articulating appendages, attitude kinematics, and the controller dynamics.  
The constraints for the OCP include the attitude, rate, and momentum states at the beginning and end of the 
maneuver, as well as peak jet torques. Additional constraints such as maximum angular excursions in each axis can 
also be specified. The degrees of freedom are the commanded vehicle attitude and rate history with respect to the 
LVLH frame. It should be noted that in general the maneuver trajectory is a function of the particular ISS 
configuration mass properties as well as the specific combination of rotary joint motions (e.g. solar arrays). 
The details of this nonlinear, constrained optimal control problem are similar to that described in Ref. 7. The 
optimal control problem is solved using the software package DIDO8. 
III. Operational Implementation 
To implement the OPM, the ground-developed trajectory commands are uploaded to the ISS. The Attitude 
Determination and Control Officer (ADCO) receives this data from the Mission Evaluation Room (MER) GN&C 
Team and uses it as input to a software tool which builds a GMT time-tagged command pair sequence for uplink to 
the Command and Control computer (C&C MDM) prior to the maneuver execution time. Since the C&C MDM 
command buffer is limited to 900 slots, the OPM is allocated 110 slots and is composed of 55 commands. 
Since the OPM is a feedforward open loop trajectory that is a function of initial states and ISS dynamics, its 
performance depends on how accurately these are known, e.g. mass properties, center of pressure, aerodynamic drag 
force, etc. Robustness analysis is performed to identify the range of uncertainty in these parameters. Similarly, when 
handing-off from the maneuver to momentum manager there will also be discrepancies in the rotational states from 
their target values. If these discrepancies are outside the robustness bounds of momentum manager startup, more 
propellant may be used to initialize the rotational states. 
IV. Flight Demonstration Description & Predicted Performance 
The first OPM operational demonstration was prior to docking of Progress (48P) to ISS, and involved a 
maneuver from +X-axis in Velocity Vector (+XVV) to –X-axis in Velocity Vector (–XVV) flight attitude.  The ISS 
GN&C would be in +XVV momentum management, then transition to USTO control to perform the maneuver and 
then transition into the –XVV momentum manager. The initial and final Local Vertical Local Horizontal (LVLH) 
attitude targets were  6.02.06 degrees and  6.08.0175 degrees (YPR order and sequence) respectively. 
 
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The initial CMG momentum in LVLH frame was  100000  ft-lbf-s and the final momentum target was  10000500 ft-lbf-s (RPY order). The maneuver duration was 5500s, with 55 uniformly spaced attitude/rate 
command pairs (updated every 100s). The expected ISS altitude was 224 nautical miles and the expected solar beta 
angle was 52deg. The atmosphere parameters were based on predicted NOAA9 environmental conditions of 
F10.7=120 (solar radio noise flux) and Ap=5 (geomagnetic index). The rotary joint operations were assumed to be 
as follows: 
 Port Solar Array Rotary Joint (SARJ): Fixed at 270 
 Starboard SARJ: Fixed at 90 
 P4,P6,S4,S6 Beta Joints: Autotrack 
 Port & Starboard Thermal Radiator Rotary Joints (TRRJs): Fixed at 45 
The second OPM demonstration was scheduled after Progress 48P had docked to ISS and involved a maneuver 
from –XVV to +XVV flight attitude.  The ISS GN&C would be in +XVV momentum management, then transition 
to USTO control to perform the maneuver and then transition into the –XVV momentum manager. The initial and 
final Local Vertical Local Horizontal (LVLH) attitude targets were  6.03.0175  degrees and  6.08.06  degrees (YPR order and sequence) respectively. The initial CMG momentum in LVLH frame 
was  100001000 ft-lbf-s and the final momentum target was  5000500  ft-lbf-s (RPY order). The 
maneuver duration was 5500s, with 55 uniformly spaced attitude/rate command pairs (updated every 100s). The 
expected ISS altitude was 224 nautical miles and the expected solar beta angle was 56deg. The atmosphere 
parameters were based on predicted NOAA9 environmental conditions of F10.7=120 (solar radio noise flux) and 
Ap=5 (geomagnetic index). The rotary joint operations were assumed to be as follows: 
 Port Solar Array Rotary Joint (SARJ): Fixed at 270 
 Starboard SARJ: Fixed at 90 
 P4,P6,S4,S6 Beta Joints: Autotrack 
 Port & Starboard Thermal Radiator Rotary Joints (TRRJs): Fixed at 45 
 
To assess the robustness of the OPM, Monte-Carlo simulations were performed using the high-fidelity Space 
Station Multi-Rigid Body Simulation (SSMRBS)10 for various parameter uncertainties. A single variable was 
perturbed at a time to identify sensitivity to individual error sources.  Results for initial state and inertia 
perturbations are given in Figure 1, which shows propellant consumed for the OPM for 500 samples. It is seen that 
the OPM can tolerate a wide range of errors, but completing the maneuver with large perturbations without 
redesigning the optimal trajectory costs more propellant. Figure 2 shows similar results for the second OPM. 
 
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Figure 1. OPM1 robust performance. 
 
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The robustness of OPM to mass properties is advantageous in the presence of ISS visiting vehicle schedule 
uncertainties since a single OPM can be used for multiple configurations. Figure 3 shows the expected visiting 
vehicles for the OPM flight demonstration, including the 48P Progress docked at the DC1 port. Since HTV3 was to 
be berthed days before the first OPM demonstration, we considered the impact of using the unmodified OPM to 
maneuver the ISS without HTV berthed. Figure 4 and Figure 5 show the OPM is robust to the absence or presence 
of HTV as well as the Progress at DC1. 
 
Figure 2. OPM2 robust performance. 
 
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Figure 3. ISS visiting vehicles. 
 
Figure 4. OPM1 robust performance to visiting vehicle uncertainty. 
 
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V. Flight Results 
The OPM flight tests took place on August 1, 2012 and surrounded the launch and docking of the 48P Progress 
vehicle. The first OPM began at GMT 214:11:55 and ended at GMT 214:13:25, rotating the ISS 180deg in 
preparation for 48P docking. Then 48P launched at GMT 214:19:35, successfully executing a new fast rendezvous 
technique to dock to ISS at GMT 215:01:18. The second OPM was then initiated at GMT 215:04:15, and ended at 
GMT 215:05:45, putting the ISS back in an orientation suitable for long-term attitude hold with momentum 
manager. Both OPMs completed successfully and used 9.7kg and 10.2kg respectively. The Mission Evaluation 
Room (MER) console views of the first OPM are shown in Figure 6Figure 8. Figure 6 shows the ISS actual attitude 
and the commands, while Figure 7 shows the total CMG momentum magnitude, the total CMG momentum vector, 
and momentum as a percent of total CMG capacity. Figure 8 shows the estimated propellant consumption 
immediately after maneuver completion. The propellant consumption values quoted earlier are based on finalized 
estimates from processed telemetry. Similarly, Figure 9Figure 11 show the attitude, CMG momentum, and 
propellant for the second OPM. 
 
Figure 5. OPM2 robust performance to visiting vehicle uncertainty. 
 
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Figure 6. OPM1 flight commanded and actual attitude. 
 
Figure 7. OPM1 flight total CMG momentum. 
 
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VI. USTO Comparison 
For comparison, the identical maneuvers from the demonstrations were simulated with USTO without using OPM. 
Flight data from the demonstrations were used to set the initial conditions. The attitude and propellant consumption 
are shown in Figure 12 for the first OPM, and Figure 13 for the second OPM. It is seen that the identical maneuver 
using USTO consumes 150kg. 
 
Figure 8. OPM1 estimated flight propellant consumption. 
 
Figure 9. OPM2 flight commanded and actual attitude. 
 
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Figure 10. OPM2 flight total CMG momentum. 
 
Figure 11. OPM2 estimated flight propellant consumption. 
 
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VII. Conclusions 
This paper presented the results of the first ever flight demonstration of the Optimal Propellant Maneuver 
(OPM). On August 1, 2012, OPM was successfully used to complete two ISS docking maneuvers which saved 93% 
propellant compared to standard methods. Additionally, OPM was designed to not require any flight software 
modifications, which was achieved by only uploading time-tagged commands. It also reduces ISS structural loads 
compared to typical ISS rotations using thruster control. This OPM was historic at several levels. From an economic 
perspective, the amount of propellant saved was approximately 280kg for an estimated value of over $9M. Feedback 
implementation is also possible, however, to implement it would require modifications to the ISS flight software. 
Moreover, the OPM flight tests established that “low-cost” rotational state transition using thrusters could be 
executed while satisfying system constraints, much like the ZPM flight tests established for momentum actuators. 
Thus the OPM and ZPM concepts can be applied to any spacecraft controlled using thrusters or momentum 
actuators respectively. 
References 
 
1Bedrossian, N., Jang, J.-W., Alaniz, A., Johnson, M., Sebelius, K., and Mesfin, Y., “International Space Station US GN&C 
Attitude Hold Controller Design for Orbiter Repair Maneuver,” AIAA Guidance, Navigation, and Control Conference, San 
Francisco, CA, Aug. 15–18, 2005, AIAA-2005-5853, pp. 1–15. 
2Liu, S.-W., and Singh, T., “Fuel/Time Optimal Control of Spacecraft Maneuvers,” Journal of Guidance, Control, and 
Dynamics, Vol. 20, No. 2, 1997, pp. 394–397. 
3Dixon, M., Edelbaum, T., Potter, J., and Vandervelde, W., “Fuel Optimal Reorientation of Axisymmetric Spacecraft,” 
Journal of Spacecraft and Rockets, Vol. 7, No. 1, 1970, pp. 1345-1351. 
 
Figure 12. Simulated USTO performance for identical maneuver as OPM1. 
 
Figure 13. Simulated USTO performance for identical maneuver as OPM2. 
 
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4Lowry, N., and Hall, R., “An Improved Three-Pulse Algorithm for Axisymmetric RCS Control,” AIAA Guidance, 
Navigation, and Control Conference and Exhibit, Austin, TX, Aug. 11-14, 2003, AIAA-2003-5329. 
5Bedrossian, N., Bhatt, S., Lammers, M., Nguyen, L., and Zhang, Y., “First Ever Flight Demonstration of Zero Propellant 
Maneuver Attitude Control Concept”, Proceedings of 2007 AIAA Guidance, Navigation, and Control Conference, Hilton Head, 
SC, Aug. 20-23, 2007, AIAA 2007-6734, pp. 1-12. 
6Bedrossian, N., Bhatt, S., Lammers, M., and Nguyen, L., “Zero Propellant Maneuver Flight Results for 180° ISS Rotation,” 
in Proceedings of the 2007 International Symposium on Space Flight Dynamics, Annapolis, MD, Sept. 24–28, 2007, NASA/CP-
2007-214158, pp. 1–10. 
7Bhatt, S., “Optimal Reorientation of Spacecraft Using Only Control Moment Gyroscopes”, Master’s Thesis, Dept. of 
Computational & Applied Mathematics, Rice Univ., Houston, TX, May, 2007. [online] URL: 
http://www.caam.rice.edu/tech_reports/2007/TR07-08.pdf. 
8Elissar, Monterey, CA.. DIDO [online] URL: http://www.elissar.biz/.  
9NOAA website, [online] URL: http://www.sec.noaa.gov/ftpdir/weekly/27DO.txt. 
10Bennett, J., "Version Baseline 3.9 Release of Space Station Multi-Rigid Body Simulation," JSC-28702Q, March, 2009. 
 


Optimal Propellant Maneuver Flight Demonstrations on ISS

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