Transcript OMNI Mission

MAXIM
Power Subsystem
Diane Yun
Vickie Moran
NASA/GSFC Code 563
[email protected]
[email protected]
301-286-0110 (IMDC)
8/19/99
Mission Description
 MAXIM is a straight-forward mission for power system design.
 The only area of concern is the uncertainty in the instrument loads; however,
the mission is not weight, surface area, or launch vehicle volume constrained
which means that larger solar arrays or batteries could be incorporated later if
needed.
 The MAXIM mission consists of two spacecraft formation flying in a flyaway orbit which is essentially heliocentric resulting in no eclipse periods.
 Each spacecraft will be maintained solar inertial (i.e. having one axis pointed
at the sun continuously).
 The “Optical” spacecraft contains the optical mirrors.
 The “Detector” spacecraft maneuvers to maintain its detector in line with the
focus of the mirrors on the “Optical” spacecraft and ~450km from the
“Optical” spacecraft.
Loads Rack-Up & Assumptions
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Optical Spacecraft--521W including 20% contingency
 259W Instrument Load
 262W Spacecraft Load
Detector Spacecraft--566W including 20% contingency
 313W Instrument Load
 253W Spacecraft Load
Both spacecraft have identical attitude control sensors and actuators including: CSS, IRU, 4 RWs, Thrusters,
and a Star Tracker.
The reaction wheels are assumed to operate at an average of 18W each except when maneuvering for new
targets when the power goes up to 80W each wheel (approximated for power system sizing as 2 hours each
day).
The Power System Electronics (PSE) is a Direct Energy Transfer (DET) architecture similar to MAP which is at
L2. Power Distribution is included in the MAP PSE.
The detector spacecraft has a low power S Band Receiver and Transmitter (20W). Both are on continuously.
The optical spacecraft has a low power S Band Receiver and Transmitter (20W). The receiver (5W) is on
continuously. The transmitter is only on for maneuvers to new targets (approximated for power system sizing as
2 hours each day).
The optical spacecraft has a X Band Receiver and Transmitter. The receiver (5W) is on continuously. The
transmitter (32W) is on ~ 1 hour each day.
The C&DH is estimated to consume equivalent power (30W) on each spacecraft and incorporate all ACE
functions.
We assume less heater power for the detector spacecraft (15W) than the optical spacecraft (30W) because of the
lower energy density.
Loads Rack-Up
Normal Mode-Duty
Detect or
Cy cle (%) Spacecraf t
Tot al Power (Watt s)
Cont ingency
Instrument Contingency (%)
Bus Contingency (%)
Tot al Power
20. 00
20. 00
Instruments
Detect or Spacecraf t
Quantum Calorimet er
Target Acquisition CCD
Station-Keeping CCD
Normal Mode-Optics Spacecraf t
566.0
521.2
52. 2
42. 1
471.7
43. 2
43. 7
434.4
261.0
216.1
100.0
100.0
100.0
100
150
11
0
0
0
100.0
0.0
61. 5
8.3
100.0
100.0
100.0
100.0
0.0
0.0
0.0
0.0
0.0
5.0
49. 0
50. 0
0.6
50
Optics Spacecraf t
Heat ers
Optics (Motors) Assumptions: 10W each
motor during adjust ; 6 mot ors max on
simultaneously; On Time=2/ 24hours
Station Keeping Int erf erometer
Pitch & Y aw I nt erf erometers
Mirror Alignment LASER
Wolt er CCD
3@200mW each
Spacecraf t Loads
PSE/ PSDU
IRU
Course Sun Sensor
Reaction W heels
Thrusters
Star Tracker
C&DH
S Band Transponder/ Transceiv er (Detect or)
X Band Receiv er
X Band Transmitter
S Band Receiver On Optics S/C
S Band Transm itt er On Optics S/C
Thermal
MAP Type PSE (90% ef f )
21W w/o Heat ers 30-39W w/Heat ers
[email protected]
4@18W av g/80W peak each (avg: 22/24 hrs; pk: 2/ 24 hrs)
On sec/day normal mode (neglected) + 1.0W Transducers
Also Does ACE Functions
Receiv er 5W on cont inuously ; Transm itt er on (15W) on cont inuously
On Receiv er; 5W on continuously f or gnd cm d link
32W On 1/24 Hrs
5W Receiv er On Alway s
15W Transmit ter On 2/ 24
Bat tery & Propulsion Heat ers
100.0
100.0
100.0
100.0
100.0
100.0
100.0
100.0
100.0
4.2
100.0
8.3
100.0
210.7
218.3
66. 0
39. 0
0.0001
92. 7
1.0
13. 0
30. 0
20. 0
0.0
0.0
0
0
15. 0
61. 0
39. 0
0.0001
92. 7
1.0
13. 0
30. 0
0.0
5
1.3
5.0
1.3
30. 0
Power System Sizing Philosophy
 The system was sized to provide energy balance on a 24 hour basis.
 There are no eclipses but we still need to carry a battery to supply power in the
launch vehicle until spacecraft sun acquisition on orbit. We minimize the solar
array area required by using the battery to supply supply transient peaks in the
load over the 24 hours.
 The solar array is sized to provide the average load and recharge the battery
over the 24 hour period.
Battery Sizing
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The peaks identified to date (for the Optics spacecraft) are:
 The optics motors during a mirror adjust. (6 motors running simultaneously at 10W each 2 hours. SA is
sized for 5W average; energy out of battery=(60W-5W)*2hrs=110Wh)
 The reaction wheels slewing the spacecraft for target acquisition. (4 wheels running at 80W each for 2
hours. SA is sized for 92.7W average; energy out of battery=(320W-92.7W)*2hrs=454.6Wh)
 The X Band Transmitter running at 32W for 1 hour each day. SA is sized for 1.3W average; energy out of
battery=(32W-1.3W)*1hr=30.7Wh
 The S Band Transmitter on the Optics S/C (15W) for target acquisition 2 hours each day. SA is sized for
1.25W average; energy out of battery=(15W-1.25W)*2hrs=27.5Wh
 Total Energy Out Of Battery Each Day=622.8Wh=22.2Ah
 ~60Ah battery provides a maximum DoD of 46% for 3 year life. 40-50% DoD max is recommended
for GEO (24 hour charge/discharge cycles).
The peaks identified to date (for the Detector spacecraft) are:
 The reaction wheels slewing the spacecraft for target acquisition. (4 wheels running at 80W each for 2
hours. SA is sized for 92.7W average; energy out of battery=(320W-92.7W)*2hrs=454.6Wh)
 Total Energy Out Of Battery Each Day=454.6Wh=16.2Ah
 Because the energy out of the battery could be much higher than calculated above due to detector
alignment with the optics spacecraft and because there is no weight issue with the launch vehicle, we have
specified the same battery size for the detector spacecraft as the optics spacecraft. Further analysis of the
power required for the detector alignment maneuvers needs to be done.
10 years from now it is reasonable to assume that rechargeable Lithium Ion batteries will be available in large
capacity sizes with energy densities of ~120Wh/kg; Battery Weight: 14kg
Solar Array Sizing--Optical Spacecraft
 The solar array is body-mounted on the half of the cylindrical surface that faces the
sun.
 The solar cells are mounted to a band around the body of the cylindrical spacecraft.
The band dimensions are 3.8m dia x 2.0m length. The projected area is 7.6m2.
 We assume that 85% of that area is available for mounting active solar cells.
 Sizing Results (assuming 6% loss in efficiency from BOL to EOL):
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Technology
Projected Area
Surface Area
Si
6.76m2
10.62m2
GaAs
4.33m2
6.80m2
Triple-Junction GaAs 3.33m2
5.24m2
*Does Not Include Substrate (included in mechanical budget)
Weight*
15.8kg
11.6kg
8.9kg
ROM Cost
1038K
2224K
2231K
Solar Array Sizing--Detector Spacecraft
 Because the body of the spacecraft is too small to support a body-mounted solar array,
the solar array is mounted to fixed, deployed panels. The panels are maintained
normal to the sunline by the spacecraft.
 The projected area is 6m2.
 We assume that 85% of that area is available for mounting active solar cells.
 Technology
Projected Area
Weight*
ROM
Cost
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7.16m2
4.58m2
3.53m2
29.2kg
19.7kg
15.1kg
700K
1499K
1503K
Si
GaAs
Triple-Junction GaAs
*Includes Substrate
Solar Array Output--Body Cells
 The solar array was sized assuming 11.1% loss in power from BOL to EOL & 90°C SA
temperatures.
 We do not have the capability to perform radiation analyses for non-Earth orbiting
spacecraft. Since weight is not a problem a thicker coverglass can be applied to the cell to
protect against radiation damage.