1、m Ob95534 0003284 BO5 = Special Copy right Notice O 1994 by the American Institute of Aeronautics and Astronautics. All rights reserved. R-060-1993 Recommended Practice Recommended Practice for Reporting Earth=to=Orbit Mission Profiles AIAA R-060- 1993 Recommended Practice for Reporting Earth- to-Or
2、bit Mis sion Profil Dennis Haas of Lockheed Missiles and Space Company, Inc. Space System Division for his inputs on orbit state parame- ters; Derek E. Lang of the Office of Com- mercial Space Transportation, Department of Transportation for keeping the standards committee informed on the activities
3、 of the US government and their regulatory con- cerns; Capt. Todd Freece of the US Air Force ESMC .for first voicing his concerns as a launch vehicle customer and stating the need for this standard; finally, the AIAA Space Launch Systems Committee on Standards chaired by Todd J. Mosher for supportin
4、g and providing comments for inclusion in this project. This document was approved by the Space Launch Systems Committee on Standards in March 1993. The AIAA Standards Technical Council (A. H. Ghovanlou, Chairman) approved the document in May 1993. V AIAA R-Ob0 93 m Ob95534 0003526 956 m AIAA R-060-
5、1993 vi 1.0 INTRODUCTION 1.1 Purpose The performance analyst, government regu- latory agencies, vehicle manufacturers, and spacecraft communities need a standard set of orbit missions and launch sites to accom- modate comparisons of launch vehicle pay- load capability. A set of common defini- tions
6、allows easy comparisons when select- ing launch configurations, setting and paying fees for launch services, and allowing mea- sured quantitative comparison between vehicles and capabilities. 1.2 scope Definitions and terms used in vehicle per- formunce comparisons are specified. Orbital profiles ar
7、e specified for generating compar- ison missions (e.g., 185 km circular orbit from Cape Canaveral). A table of current launch site latitudes is provided. A perfor- mance chart format giving information of a system payload capability is also described. 2.0 LAUNCH-TO-ORBIT MISSIONS Launch-to-orbit mis
8、sion profiles are used when generating performance statements re- garding the payload capability of a launch vehicle which carries a payload from the earths surface into orbit. For purposes of comparison, payload capability is defined as payload separated mass. These profiles form a suite of cases t
9、hat should be used for comparing the capabili- ties of a launch vehicle system. A subset of all possible launch points are used to keep the number of cases manageable and to provide good point-to-point comparison. When generating the performance quotation, a rotating earth should be used. Geophysica
10、l constants and atmospheric model should be referenced by the generator of the orbital cases. For example, the WGS-84 earth model and Jacchia 57 1 reference atmosphere AMA R-060-1993 are recommended for specifying these characteristics. The launch-to-orbit profiles are stated be- low. Each final orb
11、it mission has a name or type, perigee altitude, apogee altitude, incli - nation. It also has an argument of perigee (if applicable) and an associated launch site latitude or initial inclination for transfer orbits. All altitudes refer to local mean sea level. 2.1 Low Earth Orbit - ER This low earth
12、 orbit mission launched due east from the Eastern Range (ER) at Cape Canaveral, Florida has the lowest altitude orbit. Perigee Altitude 185km (100nmi) Apogee Altitude 185km (100nmi) Inclination 28.5 O Launch Latitude 28.5 O 2.2 Low Earth Orbit - DOT This low earth orbit mission launched from the Eas
13、tern Range (ER) at Cape Canaveral, Florida is the same mission specified by the US. Department of Transportation for com- puting the payload capability of a launch vehicle: the maximum payload to this orbit is used for calculating launch fees. Perigee Altitude Apogee Altitude Inclination 28.5 O Laun
14、ch Latitude 28.5 O 277.8 km (150 nmi) 277.8 km (150 nmi) 2.3 Low Earth Orbit - WR The low earth polar orbit mission launched in a southerly direction out of the Western Range (WR) at the Vandenberg Air Force Base, California. Perigee Altitude Apogee Altitude Inclination 90.0 O Launch Latitude 34.7 O
15、 185 km (100 nmi) 185 km (100 nmi) AIAA R-Ob0 93 9 Ob95534 0001528 729 AIAA R-060-1993 2.4 Molniya Orbit This is a 12-hour period eccentric orbit at the critical inclination. A satellite will linger near two points in the northern hemisphere, separated 180 O in longitude, each day. Perigee Altitude
16、740 km (400 nmi) Apogee Altitude 39612 km (21389 iiini) Argument of Perigee 270 O Final Inclination 63.4 O 2.5 Geostationary Transfer Orbit The process of launching a satellite from the earths surface into geostationary transfer or- bit is usually accomplished in two steps. First, a launch occurs in
17、to a trajectory hav- ing an apogee altitude of approximately 185 kin, with trajectory inclination of 28.5 O. Then through an orbital maneuver the payload achieves the final parameters given below. The orbit is further constrained to have the apogee occur over the equator. Perigee Altitude 185km (1OO
18、nmi) Apogee Altitude 35786 kin (19323 nrni) Final Inclination 26.5 O Initial Inclination 28.5 O Launch Latitude 28.5 O Argument of Perigee O O or 180 O 2.6 Geostationary Orbit The orbital transfer condition for the geosta- tionary orbit is the circularizing of the trans- fer orbit along with changin
19、g the inclination of the orbit. This orbit has a 24-hour period which places the satellite over the equator at a desired longitude. It is commonly used for conununication satellites. Perigee Altitude Apogee Altitude Final Inclination 0.0 O Initial Inclination 26.5 O Launch Latitude 28.5 O 35786 km (
20、19323 nmi) 35786 km (19323 nmi) 2.7 Geostationary Transfer Orbit, Kourou The process of launching a satellite from the earths surface into geostationary transfer orbit from Kourou, French Guiana is usually accomplished in one step. A launch occurs into a ballistic-type trajectory having an apogee al
21、titude of approximately 200 km, with trajectory inclination of 7.0 O. The orbit is further constrained to have the apogee occurring over the equator. Perigee Altitude 200 km (108 nmi) Apogee Altitude 35786 km (19323 nmi) Inclination 7.0 O Launch Latitude 5.2 O Argument of Perigee 178 O 2.8 Geostatio
22、nary Orbit, Kourou The orbital transfer conditions for the geostationary orbit is the circularizing of the transfer orbit along with changing the inclination of the orbit. These particular parameters are used for systems, launched from Kourou launch site. Perigee Altitude 35786 km (19323 mni) Apogee
23、 Altitude 35786 km (19323 nmi) Final Inclination 0.0 O Initial Inclination 7.0 O Launch Latitude 5.2 O 2.9 Geosynchronous Orbit This is a %hour period orbit for a satellite. It is used by systems that require a 24-hour period, but need not be stationary over the equator. Perigee Altitude Apogee Alti
24、tude Inclination 45.0 O Launch Latitude 28.5 O Argument of Perigee 270 O 23 138 km (12494 nmi) 48435 km (26153 nmi) 2 2.10 Sun-Synchronous Orbit This mission is launched in a southwesterly direction from the Western Range (WR) at Vandenberg Air Force Base, California. The inclination and circular al
25、titude of this orbit are selected such that the regression of the orbit plane due to Earth oblateness matches the rate at which the Earth orbits the Sun. Perigee Altitude Apogee Altitude inclination 98.6 O Launch Latitude 34.7 O 800 km (432 nmi) 800 kni (432 nmi) 3.0 LAUNCH LATITUDES A list of latit
26、udes for several launch com- plexes in the world is supplied as a reference for the developer of performance quotes. Sun hhrco Kourou Sri harikota Xi Chang Canaveral Tanegashima Negev Kagoshima Vandenberg Wallops Jiuquan Ty uratam Kapustin Yar Plesetsk 2.9“ S 5.2“ N 13.9“ N 28.2“N 28.5“ N 30.2“ N 31
27、.0“ N 31.2“ N 34.7“ N 37.9“ N 40.7“ N 45.6“ N 48.4“ N 62.8“ N Kenya (USA) French Guyana India China USA Japan Israel Japan USA USA China Russia Russia Russia ATAA R-060-1993 4.0 MISSION PERFORMANCE CHART A format for displaying the payload-to-orbit performance of a launch system is described. The pe
28、rformance chart displays the maximum payload, defined as weight above the interface between the vehicle and the payload, to specified orbits for the useful span of inclinations. When developing these charts for a launch vehicle, the minimum inclination angle line shown should be equal to the minimum
29、 latitude from which the launch system would ever be launched. The typical series of inclinations that will be used are, 7.0, 28.5“, 45.0“, 55.0“, 63.4“, 90.0, and 99.0“. The sample performance chart shown in ,Figure 1 has three distinct regions; ballistic trajectories (orbits that intersect the ear
30、th), direct ascent elliptical orbits, and direct ascent circular orbits. The three lines for each inclination all intersect at the 185 km altitude point. The ballistic trajectories (earth-intersecting or fast-decaying orbits) have apogee altitudes of 185 km and perigee altitudes of 185 km or less. D
31、irect ascent elliptic orbits will have perigee altitude of 185 km and apogee altitudes greater than 185 km. The direct ascent circular orbits will have altitudes greater than or equal to 185 km. Other performance considerations should also be indicated on the chart. For example, identify the range i
32、mposed launch site incli- nation constraints. Scales should be adjusted for appropriate presentation. 3 AIAA R-Ob0 93 0695534 0001530 387 ATAA R-060-1993 I I I Direct Ascent Elliptical Orbit, Ballistic Trajectories Perigee Altitude = 185 km Apogee Altitude = 185 km - Direct Ascent Circular Orbits .
33、. L. .- . . . . , L, 1. I h 1- 1- i- ., I , + -2000.0 o .o 2000 .o 4000.0 6000 .o 8000.0 100 Payload Weight (kg) 0 .O Figure 1 Booster Performance Chart AIAA R-Ob0 93 m Ob95534 OOOL53L 213 m American Institute of Aeronautics and Astronautics The Aerospace Center 370 LEnfant Promenade, SW Washington, DC 20024-251 8 ISBN 1 -563471O53-5
