AIAA S-122-2007 Electrical Power Systems for Unmanned Spacecraft《无人飞船的电源系统》.pdf

1、 Standard AIAA S-122-2007 Electrical Power Systems for Unmanned Spacecraft AIAA standards are copyrighted by the American Institute of Aeronautics and Astronautics (AIAA), 1801 Alexander Bell Drive, Reston, VA 20191-4344 USA. All rights reserved. AIAA grants you a license as follows: The right to do

2、wnload an electronic file of this AIAA standard for storage on one computer for purposes of viewing, and/or printing one copy of the AIAA standard for individual use. Neither the electronic file nor the hard copy print may be reproduced in any way. In addition, the electronic file may not be distrib

3、uted elsewhere over computer networks or otherwise. The hard copy print may only be distributed to other employees for their internal use within your organization. AIAA S-122-2007 Standard Electrical Power Systems for Unmanned Spacecraft Sponsored by American Institute of Aeronautics and Astronautic

4、s Approved 5 January 2007 Abstract This document, when followed in its entirety, will yield a robust EPS design suitable for very high-reliability space missions. This document specifies general design practices and sets minimum verification and validation requirements for power systems of unmanned

5、spacecraft. The focus of the document is on earth orbiting satellites using traditional photovoltaic/battery power, but does not exclude other primary power generation and storage methods. This document does not address specific launch vehicle requirements however much of the design philosophy used

6、here is applicable to launch vehicle power systems. AIAA S-122-2007 ii Library of Congress Cataloging-in-Publication Data Standard electrical power systems for unmanned spacecraft / sponsored by American Institute of Aeronautics and Astronautics. p. cm. ISBN 1-56347-913-3 (hardcopy) - ISBN 1-56347-9

7、14-1 (electronic) 1. Space vehicles-Auxiliary power supply 2. Artificial satellites-Power supply. I. American Institute of Aeronautics and Astronautics. TL1100.S83 2006 629.47445-dc22 2006038837 Published by American Institute of Aeronautics and Astronautics 1801 Alexander Bell Drive, Reston, VA 201

8、91 Copyright 2007 American Institute of Aeronautics and Astronautics All rights reserved No part of this publication may be reproduced in any form, in an electronic retrieval system or otherwise, without prior written permission of the publisher. Printed in the United States of America AIAA S-122-20

9、07 iii Contents Foreword .vii 1 Scope 1 2 Tailoring.1 3 Applicable Documents1 4 Vocabulary 2 4.1 Acronyms and Abbreviated Terms 2 4.2 Terms and Definitions.4 5 Purpose of EPS 10 5.1 Functional Description 10 5.1.1 Power Generation/Energy Conversion 10 5.1.2 Energy Storage .10 5.1.3 Power Management

10、and Distribution (PMAD) 11 5.1.4 Flight Software 11 5.1.5 Harness .11 5.2 Functional and Performance Requirements12 5.2.1 Orbital Profile 12 5.2.2 Mission Life .13 5.2.3 Mission Phases.13 5.2.4 EPS Architecture.13 5.2.5 Power Generation.14 5.2.6 Energy Storage .16 5.2.7 Power Management19 5.2.8 Powe

11、r Distribution.20 5.2.9 Grounding and Bonding22 5.2.10 Energy Management 22 5.2.11 Telemetry and Command.24 5.2.12 Power Quality25 5.2.13 EMI/EMC .27 5.2.14 Fault Detection, Isolation, and Recovery.27 5.3 Design and Construction Requirements28 5.3.1 Parts, Materials and Processes (PMP) .28 5.3.2 Pro

12、duct Markings29 5.3.3 Manufacturing Management 29 5.3.4 Interchangeability a logical load group may be a set of loads to be turned off in safe-hold or survival mode. Load Margin power margin divided by the contingent load power, often stated as a percentage NOTE Load margin represents the additional

13、 percentage of load power the source could provide and still recharge the batteries. Lock-up (also called latch-up) condition wherein the solar array is operated at a point well below its maximum power point such that the solar array output power is insufficient to power the load and recharge the ba

14、ttery NOTE Lock-up arises in EPS architectures that have multiple stable operating points determined by combined solar array, battery charge controls, battery, and load characteristics. When lock-up occurs, system operation is at an operating point below the solar array maximum power point, resultin

15、g in insufficient battery recharging and, possibly, some battery discharge. AIAA S-122-2007 7 Losses distribution losses and energy storage system inefficiency losses associated with supporting the spacecraft electrical loads Main Bus main conductors to which the sources are attached prior to any br

16、anching NOTE The impedance of this bus is common to all loads. Mission Life the contractually required period of time over which the spacecraft must meet all of its performance requirements NOTE Mission life includes transfer orbit, orbit-raising, on-orbit, and de-orbit modes. EPS required life or a

17、n EPS component required life can be longer than the spacecraft mission life requirement. A design life can be longer than a required life. Multipoint Ground allows use of structure as a low-impedance return path for currents NOTE Care must be taken to avoid large DC currents that can interfere with

18、 low-level circuitry. Multipoint grounding offers some advantages for high-frequency subsystems. Nonessential Loads loads that can be powered off without adversely affecting the minimum controllability and commandability of the spacecraft Normal Operation range of operational states of the space veh

19、icle that exist or occur by design and in which the vehicle spends the vast majority of its time, according to the expectations of the mission designers and planners Operational States all foreseeable and intentional combinations of states, modes, or conditions within the EPS hardware and software P

20、ayloads self-contained instrument, sensor, or device that fulfills some mission objective Periodic and Random Deviation (PARD) instantaneous time-domain deviation of a voltage about a nominal value as observed on an oscilloscope with a specified bandwidth and stated as a peak-to-peak or root-mean-sq

21、uare value Power Budget method of accounting for the spacecrafts electrical loads and losses associated with these loads NOTE Often, loads are expressed in terms of orbital average power. It may also contain, as required by the energy balance analysis, information on the time behavior of loads. This

22、 could be expressed in terms of an orbital duty cycle or a detailed, moment-by-moment power prediction. A power budget usually contains multiple load predictions corresponding to different mission modes or DRCs. Power Control all hardware and software used to control and steer electrical power from

23、the power generating element, to and from the energy storage subsystem, and to the power distribution subsystem Power Generation all equipment involved in the generation of DC power for use by the loads and for charging the energy storage devices AIAA S-122-2007 8 NOTE Solar arrays are the most comm

24、on technology for spacecraft power generation; other technologies, such as thermoelectric devices, fall into the category. Power Interface Loop Gain ratio of the source impedance to the load impedance NOTE The loop gain is a complex function of frequency and includes both passive elements and active

25、 contributors such as the negative impedances of DC-converters. Power Margin difference between contingent source power and contingent load power NOTE it is the margin between the lowest possible array output power and the highest possible load power. Power Quality acceptability of the time-domain v

26、ariation in bus voltage induced by the periodic and aperiodic currents flowing to and from the loads and to self-generated currents and voltages from the EPS equipment itself and by the GSE during ground testing Rated (or Nameplate) Battery Capacity minimum capacity that is based on charging the bat

27、tery at a predefined rate of charge and temperature and discharging it at the same temperature; a predefined rate of discharge and end of discharge voltage Reconditioning act of restoring lost capacity to a battery through a process of deep discharging and recharging per some prescribed charging reg

28、imen Regulated Buses one whose voltage is controlled by means of a closed-loop negative feedback control scheme NOTE A sunlight-regulated bus is regulated during insolation and unregulated during eclipse. Unregulated Buses one whose voltage is approximately the same as the battery voltage, minus har

29、ness and switching losses Safe-hold or Survival Mode state entered into by the spacecraft (typically the result of an on-board fault detection response) that maximizes the probability that the spacecraft will survive until mission operators can restore normal operations NOTE Transition to this state

30、 from normal operations typically includes the shedding of non-essential loads, and the use of redundant and/or back-up essential loads. Single-point Failure single component, wiring, or connector failure, software glitch, or computer failure that results in the permanent loss of the spacecrafts abi

31、lity to perform its primary mission for the intended design life span Single-Point Ground (SPG) a commonly-used type of voltage reference scheme (VRS) in which DC return currents from the loads or ground planes are carried via low-impedance conductors back to a single grounding point NOTE Use of str

32、ucture for DC return currents is prohibited for an SPG-type VRS. Source Power Margin power margin divided by contingent source power Spikes narrow impulse-like voltage waveforms that are produced by switching or fault-clearing events AIAA S-122-2007 9 NOTE Spikes are generally measured in terms of t

33、heir volt-second impulse strength and peak voltage amplitude. State of Charge (SOC) ratio of the number of Ah present in a battery (under specific operating conditions) to the rated capacity C(Ah) of the battery, times 100 Technology Readiness Level (TRL) numerical ranking system for assessing the m

34、aturity of and relative risk of using a particular technology NOTE For solar arrays employing crystalline solar cells, for the purposes of this document, TRL levels have been defined as follows: 9 Solar array has flown successfully on its intended mission 8 Solar array flight qualified on the ground

35、 or is flying successfully as primary power 7 Solar array prototype is flying successfully 6 Solar array has passed life-cycle and qualification test at the panel/coupon level for mission conditions 5 Solar cell/Coverglass-Interconnect-Cell (CIC) is space qualified 4 Solar cell/CIC is characterized

36、at prototype level 3 Solar cell prototype has been fabricated 2 Solar cell design and modeling completed. Subcell components demonstrated 1 Solar cell conceptual design formulated The final determination of TRL may vary somewhat from the levels defined above, depending on mission duration and critic

37、ality, orbit, and other factors. The final determination of TRL is up to the procuring authority for a particular program. Test String (I/V Measurement) at least three solar cell assemblies, electrically connected in series, with end termination tabs NOTE The test strings will be manufactured at the

38、 same time as the power generating strings. They will incorporate the same parts, materials and processes used to manufacture the power generating strings, and will be bonded to the flight panels using the same materials and processes used to bond the power generating strings. Thermal Loads dissipat

39、ive heaters used for temperature control of the spacecraft components Transients bus voltage time-domain response due to an aperiodic event or due to a periodic low-frequency train of events Virtual Cell group of cells wired in parallel Voltage Reference Point a point, usually on structure, used as

40、the reference from which other system voltage levels are calculated NOTE On an unregulated bus it is at the output of the box, which establishes the main power bus. On a regulated bus it is the output of the main bus regulator. AIAA S-122-2007 10 Voltage Reference Scheme all wiring, structures, and

41、connections that determine the return current paths in the EPS, designed to avoid interference between loads and to meet EMC requirements EXAMPLE Single Point Ground and structure-return are two examples of VRS Voltage Ripple cyclic variation of voltage about the mean level of the DC voltage during

42、steady-state operation of the EPS NOTE The ripple generally contains multiple frequency components as well as small spikes outside the average envelope of the ripple. Overall ripple is measured in RMS or peak-to-peak volts, while the spikes are generally measured in terms of their volt-second impuls

43、e strength and peak voltage amplitude. 5 Purpose of EPS The electrical power system (EPS) of a space vehicle shall be designed to ensure the reliable delivery of electrical power compatible with all loads under all foreseeable operational states and environments, during all mission phases, and over

44、the intended design life of the loads and of the space vehicle. 5.1 Functional Description Electrical power is used by all active spacecraft systems and equipment for their operation. Direct Current (DC) electrical power at different bus voltages is normally utilized for this purpose. All DC EPS inc

45、lude power generation, energy storage, power management, and distribution as well as protection devices and load shedding controls. The EPS of a spacecraft is comprised of the hardware and software used to generate, store, condition, and distribute power as required by the loads; it performs these f

46、unctions throughout all mission phases in the presence of all environments encountered. The sizing and design of all power system components take into account the spacecraft orbit and location, mission phases and durations, modes of orbit transfer, natural and man-made environmental effects, and oth

47、er mission-specific requirements such as orbital relocation and solar shadowing events. 5.1.1 Power Generation/Energy Conversion Power generation involves all equipment used in the generation of DC power for use by the loads and for charging the energy storage devices. Electric power can be generate

48、d by using solar photovoltaic, solar thermal, or nuclear energy conversion devices, depending upon the load power requirements and mission orbit and duration. The DC power systems specified in this standard assume arrays of photovoltaic solar cells which convert sunlight into useful electric power.

49、The solar arrays are sized to satisfy average power demand (including battery recharge power) during operational life in sunlight. Further, the solar arrays are sized to satisfy the power contingency requirements from beginning to end of life, including losses of power due to statistically expected performance degradation and solar cell failures. 5.1.2 Energy Storage Energy storage devices store some of the energy generated by the power generation element for use in powering the loads during eclipse periods when the output of the power generation element is insufficient