标准编号:	GJB 3862-1999
中文名称:	卫星与运载火箭接口
英文名称:	Interfaces of satellite and launch vehicle
行业分类:	GJB    国家军用行业标准
发布机构:	中国人民解放军总装备部
发布日期:	1999-08-03
实施日期:	2000-1-1
标准状态:	现行
翻译语言:	英文
文件格式:	PDF
中文字符数:	40000 字
翻译价格(元):	5000.00 元
交付时间:	付款后 1 个工作日

Codeofchina.com is in charge of this English translation. In case of any doubt about the English translation, the Chinese original shall be considered authoritative.





Interfaces of satellite and launch vehicle



1 Scope



1.1 Subject content



This standard specifies the corresponding parameters, requirements and verification method (analysis or test) for interfaces of satellite and launch vehicle, as well as the requirements for check and joint operation of interface of satellite and launch vehicle in launch site.



1.2 Application scope



This standard is applicable to the determination, verification and check of the interface relationship between various satellites and different launch vehicles, and may serve as a reference for the interface relationship between other spacecrafts and launch vehicles.



2 Normative references



GJB 151A-97 Electromagnetic emission and susceptibility requirements for military equipment and subsystems

GJB 421A-97 Satellite terminology

GJB 1028-90 Satellite coordinate system

GJB 1547-92 Technical requirements of satellite for launch vehicle



3 Definitions



3.1 Terms



The terms given in GJB 421A and the followings apply.



3.1.1 satellite system

system including satellite platform, payload and all items provided by satellite manufacturer to support launching



3.1.2 launch vehicle system

system including launch vehicle and launch services related to launch vehicle as provided by launch service contractor and its subcontractors



3.1.3 separation plane of satellite and launch vehicle

plane where the launch vehicle separates from the satellite



3.1.4 mating plane

mechanical connection plane between satellite and launch vehicle



3.1.5 payload adapter

structure connecting satellite with launch vehicle and unlocking device for connection of satellite and launch vehicle



3.1.6 usable volume

maximum volume envelope available for satellite in payload fairing of launch vehicle



3.2 Abbreviations



3.2.1 CDR command destruct receiver



3.2.2 EGSE electric ground support equipment



3.2.3 EMC electromagnetic compatibility



3.2.4 GSE ground support equipment



3.2.5 GTO geosynchronous transfer orbit



3.2.6 PLA payload adapter



3.2.7 PSD power spectral density



3.2.8 RF radio frequency



3.2.9 RT radar transponder



3.2.10 SPL sum acoustic pressure level



3.2.11 SSO sun synchronous orbit



3.2.12 TM telemetry



4 General requirements



There is no provision in this clause.



5 Detailed requirements



5.1 Mechanical interface



5.1.1 Mechanical interface state



The satellite is connected with the launch vehicle through the payload adapter (PLA).



a. When the launch vehicle manufacturer provides PLA, it shall provide the unlocking device for connection of satellite and launch vehicle simultaneously (see 5.1.4.1);



b. When the satellite manufacturer provides the unlocking device for connection of satellite and launch vehicle, it shall also provide the interface for the mating plane between satellite and launch vehicle simultaneously (see 5.1.4.2).



5.1.2 Fundamental frequency of satellite



The longitudinal and transverse fundamental frequencies of satellites shall generally not be lower than the values specified by the launch vehicle manufacturer. When lower than such values, it shall be coordinated with the launch vehicle manufacturer and confirmed after further coupling analysis.



5.1.3 Usable volume



The satellites shall adapt to the limitation of usable volume proposed by the launch vehicle manufacturer to avoid hardware collision. When the shape of the satellite partially exceeds the allowable usable volume, it must be coordinated with the launch vehicle manufacturer and confirmed after gap analysis.



See Annex A (reference) for the example of usable volume of launch vehicle.



5.1.4 PLA interface



5.1.4.1 When PLA is provided by the launch vehicle manufacturer, the interface of mating plane is determined by the following aspects:



a. Coordinate system and relative angular orientation of satellite and PLA, in which the coordinate system shall be selected according to GJB 1028.



b. Mechanical state:



Type, quantity and dimension of connecting devices (bolts and nuts), and the positions of connecting holes, dowel pins or locating slots;



Material type and characteristics;



Surface coating;



Roughness;



Flatness;



Stiffness.



c. Strap:



Geometric dimension;



Material;



Roughness;



Pretightening force (N).



Typical nominal diameters of PLA mating plane: φ937mm, φ1,194mm and φ1,497mm.



5.1.4.2 When PLA is provided by the satellite manufacturer, the interface of mating plane is determined by the following aspects:



a. Coordinate system and relative angular orientation of launch vehicle and PLA, in which the coordinate system shall be selected according to GJB 1028;



b. Mechanical state:



Type, quantity and dimension of connecting devices (bolts and nuts), and the positions of connecting holes, dowel pins or locating slots;



Material type and characteristics;



Surface coating;



Roughness;



Flatness;



Stiffness.



See Annex B (reference) for the example of PLA interface dimension.



5.1.5 Separating electric connector



The satellite manufacturer shall provide electric connectors for power supply and signal. The electrical interface of satellite and launch vehicle is completed by separating electric connectors, which are of an even number and installed symmetrically.



The following mechanical characteristics of separating electric connectors shall be specified:



a. model;



b. quantity;



c. installation position and mechanical interface;



d. plug-in and plug-out force (N);



e. mechanical separation force (N);



f. plug-in and plug-out lifetime;



g. others.



5.1.6 Separation actuator



The separation actuator provides the required energy for the separation of satellite and launch vehicle, and generally the forms of springs and retro-rockets are adopted.



5.1.6.1 Separating spring and ejector rod



The following characteristics of spring or ejector rod shall be specified:



a. quantity;



b. position;



c. normal stroke (mm);



d. compression stroke (mm);



e. maximum jacking force (N)



f. unit energy of spring or ejector rod (J).



5.1.6.2 Retro-rocket



The launch vehicle manufacturer shall provide the following characteristics of the retro-rocket:



a. quantity;



b. installation position;



c. relative speed of satellite and launch vehicle separation (m/s);



d. total impulse of launch vehicle (N·s);



e. pollution assessment.



5.1.7 Microswitch;



The satellite manufacturer shall make clear to the launch vehicle manufacturer the following characteristics of microswitch:



a. model;



b. quantity;



c. installation position and mechanical interface.



5.1.8 Operation window and microwave penetrating window



The launch vehicle manufacturer shall provide the satellite manufacturer with the required operation window and microwave penetrating window. The dimension and position of the specific window required by the satellite manufacturer shall be determined by the parties through coordination.



5.2 Electrical interface



5.2.1 Separating electric connector



The following characteristics of separating electric connectors between satellite and launch vehicle shall be specified:



a. model of connector;



b. number of cores (contacts) available to the user;



c. isolation between satellite power supply and ignition circuit of pyrotechnic device;



d. electrical performance;



e. shielding requirements;



f. locking mark.



5.2.2 Umbilical link



The umbilical link between satellite and electric ground support equipment (EGSE) is characterized as follows:



a. Quantity and form of links.



b. Limit parameters:



maximum voltage (V);



maximum current (A);



power (W);



maximum resistance or voltage drop of a single channel (Ω or V).



c. Working constraints:



quantity of functions;



function types.



d. Connector contact current.



e. Conformity confirmation:



end-end resistance (Ω);



line-ground insulation;



line-line insulation.



5.2.3 Special electrical commands for satellite



Special electrical commands for satellite, including standard commands and standby commands, are generated by launch vehicles for satellite-specific use. A list of commands shall be listed and their related characteristics shall be explained.



5.2.3.1 Pyrotechnic device command



The circuits related to the pyrotechnic device command shall be generally explained, and the corresponding schematic diagram for circuit and the following main characteristics shall be given:



a. number of commands;



b. command voltage (V);



c. command pulse width;



d. output insulation (Ω);



e. load current of pyrotechnic device (A);



f. interval between two commands;



g. insulation between wire and structure;



h. protection of satellite equipment;



i. restrictions for use on satellites.



5.2.3.2 Monitoring demonstration loop command



The circuits related to the monitoring demonstration loop command shall be generally explained, and the corresponding schematic diagram for circuit and the following main electrical characteristics shall be given:



a. number of available commands on ground or in flight;



b. command voltage (V);



c. command duration;



d. resistance (Ω);



e. on-board circuit insulation;



f. protection measures on satellites;



g. restrictions for use on satellites (maximum voltage and current).



5.2.3.3 Other electrical commands



The circuits related to such commands shall be generally explained, and the corresponding schematic diagram for circuit and the following main characteristics shall be given:



a. number of available commands on ground or in flight;



b. output voltage (V) and load current (A) of pyrotechnic device;



c. protection measures on satellites;



d. insulation between circuit and structure;



e. electromagnetic compatibility (EMC) shall meet the requirements of GJB 151A;



f. restrictions for use on satellites.



5.2.4 Separation state transmission



The separation state signals are transmitted by the launch vehicle telemetry system, and the main ways available for separation state transmission are as follows:



a. microswitch;



b. disconnection;



c. monitoring loop;



d. others.



5.2.5 TM in flight



The launch vehicle manufacturer shall explain the types of satellite data obtained from the flight TM of the launch vehicle, mainly including:



a. mechanical and thermal environment data at the interface between satellite and launch vehicle;



b. specific internal measurement parameters required for satellite.



5.2.6 Power supply



General description of circuits related to power supply shall be given, and corresponding schematic diagram for circuit, main electrical characteristics of power supply and their tolerances shall be given:



a. voltage (V);



b. current (A);



c. insulation on satellites;



d. protection measures on satellites;



e. EMC.



5.2.7 Continuity of ground potential



The electrical potential continuity of satellite to ground shall be explained:



a. position of satellite reference point (multi-points may be available);



b. maximum resistance between satellite metal parts and the nearest reference point;



c. maximum resistance of separation plane of satellite and launch vehicle.



5.3 RF/electromagnetic interface



Radiation from satellites, launch vehicles and launch sites shall be evaluated together to determine compatibility and possible limitations. The methods for verifying these interfaces are specified in 5.6 and 5.7.



5.3.1 RF telemetry (TM) and command link



The launch vehicle manufacturer shall provide TM and command link for satellite TM and command antenna and satellite system test equipment through electric connectors.



The establishment and use of this link will be realized through the RF window (i.e., microwave penetrating window) related to satellite fairing, receiving antenna or other equivalent methods. Except that the RF transmission is limited due to the operation of the launch vehicle, the link shall be available from the end of satellite installation and enclosing on launch vehicle to the launch. The TM and command frequency of the satellite must not overlap with the frequency of the launch vehicle.



The launch vehicle manufacturer shall explain the following characteristics of the launch vehicle:



a. TM:



bandwidth;



frequency;



transmitter power.



b. CDR:



bandwidth;



frequency.



c. RT:



bandwidth;



frequency;



peak power.



d. Characteristics of RF window:



material;



location (including orientation, position and dimension);



frequency range insertion loss.



5.3.2 EMC of satellite and launch vehicle



a. The satellite shall be compatible with the electromagnetic field generated by the launch vehicle, see 5.5.6.1 for the electromagnetic environment of the launch vehicle;



b. The launch vehicle shall be compatible with the electromagnetic field generated by the satellite, see 5.5.6.2 for the electromagnetic environment of the satellite.



5.4 Mission performance



This provision specifies the performance parameters of the launch vehicle to ensure that the launch vehicle can inject the satellite into the predetermined orbit.



5.4.1 General orbit



5.4.1.1 Performance



The launch vehicle’s performance of launching satellites into general orbit may be expressed by the graphical relationship between launch capability and orbit parameters (hP, ha, i and ω). See Annex C (reference), Figures C1~C3 for examples of launch capability for launching satellites into general orbit.



5.4.1.2 Injection accuracy of launch vehicle



Standard deviation (1σ) or maximum deviation (3σ) of general orbit parameters shall be given:



a. orbit inclination deviation Δi (°);



b. altitude deviation of perigee ΔhP (km);



c. altitude deviation of apogee Δha (km);



d. argument deviation of perigee Δω (°);



e. right ascension deviation of ascending node ΔΩ (°).



See Annex C, Table C1 for the example of injection accuracy of satellite (3σ).



5.4.1.3 Launch window



Any constraints of the launch vehicle and launch site on the launch window shall be specified.



5.4.2 Geosynchronous transfer orbit (GTO)



5.4.2.1 Performance



5.4.2.1.1 The geosynchronous transfer orbit (GTO) shall be determined according to the parameters of launching satellites into orbit by launch vehicles:



a. orbit inclination i (°);



b. altitude of perigee hp (km);



c. altitude of apogee ha (km);



d. argument of perigee ω (°);



e. right ascension of ascending node Ω (°).



Note: Semi-major axis a and orbital eccentricity e may be derived from the relationship between several orbit elements.



5.4.2.1.2 The performance of launch vehicle in geosynchronous transfer orbit may be expressed in the following two cases:



a. if standard PLA is adopted, the launch capability is the mass of single, double or multiple satellites launched;



b. if non-standard PLA is adopted, the launch capability is the mass of the satellite and PLA.



5.4.2.1.3 The example for the effect of altitude deviation of apogee on launch capability of satellite launched in geosynchronous transfer orbit (i=28.5°) is shown in Figure C4 of Annex C. The example for relationship between launch capability and orbit inclination of satellite launched in geosynchronous transfer orbit is shown in Figure C5 of Annex C. Detailed performance calculation of subsynchronous and supersynchronous geosynchronous transfer orbits shall be given.



5.4.2.2 Injection accuracy of launch vehicle



The injection accuracy of launch vehicle for launching satellite in geosynchronous transfer orbit shall be specified according to the parameters of geosynchronous transfer orbit (GTO).



The standard deviation (1σ) or maximum deviation (3σ) of orbit parameters shall be given. The covariance matrix of related parameters or a set of equivalent launch vehicle state parameters shall be provided as follows:



a. orbit inclination deviation Δi (°);



b. altitude deviation of perigee Δhp (km);



c. altitude deviation of apogee Δha (km);



d. argument deviation of perigee Δω (°);



e. right ascension deviation of ascending node ΔΩ (°).



5.4.2.3 Launch window



The constraints on the launch window of launch vehicle or launch site, if any, shall be expressed as a part of the standard launch window.



5.4.3 Sun synchronous orbit (SSO)



5.4.3.1 Performance



The performance of launch vehicle for satellite in sun synchronous orbit is expressed by the graphical relationship between launch capability and orbit altitude, with the example shown in Figure C6 of Annex C.



5.4.3.2 Injection accuracy of launch vehicle



The standard deviation (1σ) or maximum deviation (3σ) of sun synchronous orbit parameters shall be given:



a. altitude deviation of perigee Δhp (km);



b. eccentricity deviation Δe;



c. orbit inclination deviation Δi (°);



d. orbit period deviation ΔT (s);



e. argument deviation of perigee Δω (°).



5.4.3.3 Launch window



The constraints on launch window of sun synchronous orbit shall be given according to the local time of satellite orbit and the performance of launch vehicle.



5.4.4 Satellite orientation and separation



5.4.4.1 General description



The general description of attitude maneuverability after launching satellites into orbit by launch vehicles shall include:



a. type and propellant (gas, hydrazine) of attitude control system;



b. typical time sequence of the launch vehicle from injection to flight mission completion according to the different requirements for various satellites (three-axis stability, spin stability and multiple separations);



c. constraints on launch vehicle.



5.4.4.2 Orientation performance



The orbital coordinate system shall be specified to determine the direction requirements for the longitudinal axis orientation of spin-stabilized satellites or the two-axis orientation of three-axis stabilized satellites:



a. if the launch vehicle has this capability, necessary data for the change of satellite direction with time (as a function of launch time) shall be provided;



b. spin performance includes maximum spin speed and its error;



c. pointing accuracy of three-axis stabilized satellite (attitude error of three axes before satellite separation) and that of spin stabilized satellite (pointing error of momentum vector after satellite separation);



d. the minimum relative separation speed provided by the separation system of launch vehicle.



See Table C2 of Annex C for the example of the initial attitude angle deviation and the initial attitude angular velocity deviation (3σ) of the satellite at the end of satellite and launch vehicle separation.



5.5 Inductive environment and limit load



5.5.1 Limit load



The launch vehicle manufacturer shall give the following limit loads.



5.5.1.1 Steady-state load



The steady-state load of the launch vehicle during flight is expressed by the acceleration at the satellite mass center or the mating plane of satellite and launch vehicle, with the example shown in Figure D1 of Annex D (reference).



5.5.1.2 Quasi-static load



The quasi-static load of the launch vehicle during flight is the algebraic sum of steady-state load and dynamic load, which is expressed by the acceleration at the satellite mass center or the mating plane of satellite and launch vehicle, with the example shown in Clause D1 of Annex D.



5.5.1.3 Maximum limit load



The maximum permissible load of PLA shall be specified.



5.5.2 Mechanical environment



5.5.2.1 Low-frequency vibration



The launch vehicle manufacturer shall give the equivalent sinusoidal vibration spectrum according to the response of sinusoidal vibration and transient vibration in the relevant frequency band on the mating plane of satellite and launch vehicle, with the examples shown in Figures D2~D3 of Annex D.



5.5.2.2 Random vibration



The launch vehicle manufacturer shall give the envelope spectrum of random vibration flying in three-axial direction, with the example shown in Figure D4 of Annex D.



5.5.2.3 Noise



The launch vehicle manufacturer shall give the flight noise spectrum in the satellite fairing or the supporting structure, with the example shown in Figure D5 of Annex D. The fill factor of satellite and satellite fairing (volume ratio of satellite to fairing) shall be also indicated.



5.5.2.4 Impact



The maximum impact spectrum at the interface of satellite and launch vehicle and related areas shall be specified, with the example shown in Figure D6 of Annex D.



5.5.2.5 Satellite mass center position limit



The position of satellite mass center shall be marked and the limit of satellite mass center position and PLA weight shall be stated, with the examples shown in Annex D, D3.



5.5.3 Thermal environment



5.5.3.1 General requirements



The thermal environment of satellite includes the following aspects:



a. the thermal environment in the transfer phase of satellite from the final assembly test plant to the launch area of the launch site;



b. the thermal environment in the phase from mating of satellite and launch vehicle to pre-launch;



c. the thermal environment in the phase from satellite launch to separation of satellite and launch vehicle.



5.5.3.2 Thermal environment of ground operation



The thermal environment of ground operation mainly includes:



a. operating ambient temperature;



b. relative humidity;



c. air velocity in the fairing.



5.5.3.3 Heat flow during flight of launch vehicle



The launch vehicle manufacturer shall give the curve of heat flow at the typical reference point of the inner surface of the fairing changing with time during flight of the launch vehicle, with the example shown in Annex D, D4; for the recoverable satellite, the curve of heat flow changing with time on the outer surface of the satellite during flight of the launch vehicle shall be given.



5.5.3.4 Heat flow at the moment of satellite fairing jettisoning



The launch vehicle manufacturer shall give the maximum heat flow value heated by free molecular flow during satellite fairing jettisoning.



5.5.3.5 Heat flow in the separation phase of satellite and launch vehicle



The launch vehicle manufacturer shall give the maximum heat flow and duration generated by the launch vehicle on the satellite in the separation phase of satellite and launch vehicle.



5.5.4 Static pressure in satellite fairing



The launch vehicle manufacturer shall give the curve of static pressure in the satellite fairing changing with time during flight of the launch vehicle, with the example shown in Figure D7 of Annex D.



5.5.5 Pollution and cleanliness



5.5.5.1 Pollution of satellite



The launch vehicle manufacturer shall give the organic and particle deposits produced by the launch vehicle material outgassing and separation system to the satellite during the following satellite operations:



a. the ground phases (transportation, pre-launch) and flight phases of the satellite in the fairing (or supporting structure);



b. the smoke plume generated on launch vehicle.

1 Scope
1.1 Subject content
1.2 Application scope
2 Normative references
3 Definitions
3.1 Terms
3.2 Abbreviations
4 General requirements
5 Detailed requirements
5.1 Mechanical interface
5.2 Electrical interface
5.3 RF/electromagnetic interface
5.4 Mission performance
5.5 Inductive environment and limit load
5.6 Verification analyses and documents
5.7 Verification test
5.8 Check and joint operation requirements for interface of satellite and launch vehicle
Annex A (Reference) Examples of fairing dimension and usable volume
Annex B (Reference) Examples of interface dimensions
Annex C (Reference) Performance examples
Annex D (Reference) Examples of environmental parameters
Annex E (Reference) Verification
Additional explanation

1 范围
1.1 主题内容
本标准规定了卫星与运载火箭(以下简称星箭)相应的接口参数和要求、接口的验证方法(分析或试验)，以及发射场星箭接口检查与联合操作要求。
1.2 适用范围
本标准适用于各类卫星与不同运载火箭之间的接口关系的确定、验证和检查，其他各类航天器与运载火箭的接口关系亦可参照使用。
2 引用文件
GJB 151A-97 军用设备和分系统电磁发射和敏感度要求
GJB 421A-97 卫星术语
GJB 1028-90 卫星坐标系
GJB 1547-92 卫星对运载火箭的技术要求
3 定义
3.1 术语
除本标准定义的术语外，其他术语见GJB 421A。
3.1.1 卫星系统 satellite system
包括卫星平台、有效载荷以及卫星承制方为支持发射工作所提供的全部项目。
3.1.2 运载火箭系统 launch vehicle system
包括运载火箭以及由发射服务承包单位及其分承包单位所提供的与运载火箭有关的发射服务。
3.1.3 星箭分离面 separation plane of satellite and launch vehicle
运载火箭与卫星发生分离的平面。
3.1.4 对接面 mating plane
卫星与运载火箭之间的机械连接面。
3.1.5 有效载荷适配器 payload adapter
连接卫星与运载火箭的结构以及星箭连接解锁装置。
3.1.6 整流罩净空间 usable volume
运载火箭有效载荷整流罩内可供卫星利用的最大容积包络。
3.2 缩写词
3.2.1 CDR command destruct receiver
自毁指令接收机
3.2.2 EGSE electric ground support equipment
地面电气支持设备
3.2.3 EMC electromagnetic compatibility
电磁兼容性
3.2.4 GSE ground support equipment
地面支持设备
3.2.5 GTO geosynchronous transfer orbit
地球同步转移轨道
3.2.6 PLA payload adapter
有效载荷适配器
3.2.7 PSD power spectral density
功率谱密度
3.2.8 RF radio frequency
射频
3.2.9 RT radar transponder
雷达应答机
3.2.10 SPL sum acoustic pressure level
总声压级
3.2.11 SSO sun synchronous orbit
太阳同步轨道
3.2.12 TM telemetry
遥测
4 一般要求
本章无条文。
5 详细要求
5.1 机械接口
5.1.1 机械接口状态
卫星通过有效载荷适配器(PLA)与运载火箭连接。
a.由运载火箭承制方提供PLA时，应同时提供星箭连接解锁装置(见5.1.4.1条)；
b.由卫星承制方提供星箭连接解锁装置时，则应同时提供星箭对接面接口(见5.1.4.2条)。
5.1.2 卫星基频
卫星的纵向和横向基频一般应不低于运载火箭承制方的规定值。当低于此值时，应与运载火箭承制方协调，并进行进一步的耦合分析后予以确认。
5.1.3 整流罩净空间
卫星应适应运载火箭承制方提出的整流罩净空间的限制，以避免硬件碰撞。当卫星的外形局部超出了允许的整流罩净空间时，必须与运载火箭承制方协调，并进行间隙分析后予以确认。
运载火箭整流罩净空间的实例见附录A(参考件)。
5.1.4 PLA接口
5.1.4.1 运载火箭承制方提供PLA时对接面接口由以下方面确定：
a.卫星和PLA的坐标系及相关角方位，其中坐标系选择应符合GJB 1028中规定。
b.机械状态：
连接装置(螺栓、螺母)的类型及数量、尺寸，连接孔、定位销或定位槽的位置；
材料类型及特性；
表面涂层；
粗糙度；
平面度；
刚度。
c.包带：
几何尺寸；
材料；
粗糙度；
预紧力(N)。
典型的PLA对接面的标称直径为：φ937mm、φ1194mm、φ1497mm。
5.1.4.2 卫星承制方提供PLA时对接面接口由以下方面确定：
a.运载火箭和PLA的坐标系及相关角方位，其中坐标系的选择应符合GJB 1028中规定；
b.机械状态：
连接装置(螺栓、螺母)的类型及数量、尺寸，连接孔、定位销或定位槽的位置；
材料类型及特性；
表面涂层；
粗糙度；
平面度；
刚度。
PLA接口尺寸实例见附录B(参考件)。
5.1.5 分离电连接器
卫星承制方应提供电源和信号的电连接器。星箭的电气接口通过分离电连接器来完成，分离电连接器成偶数对称安装。
应明确分离电连接器的下列机械特性：
a.型号；
b.数量；
c.安装位置及机械接口；
d.插拔力(N)；
e.机械分离力(N)；
f.插拔寿命；
g.其他。
5.1.6 分离执行机构
分离执行机构为星箭分离提供所需的能量，一般采用弹簧、反推火箭等形式。
5.1.6.1 分离弹簧和顶杆
应明确弹簧或顶杆的下列特性：
a.数量；
b.位置；
c.正常行程(mm)；
d.压缩行程(mm)；
e.最大顶出力(N)
f.弹簧或顶杆单位能量(J)。
5.1.6.2 反推火箭
运载火箭承制方应提供反推火箭的下列特性：
a.数量；
b.安装位置；
c.星箭分离相对速度(m/s)；
d.运载火箭总冲(N·s)；
e.污染评估。
5.1.7 微动开关
卫星承制方应向运载火箭承制方明确微动开关的下列特性：
a.型号；
b.数量；
c.安装位置及机械接口。
5.1.8 操作窗口和透波窗口
运载火箭承制方应向卫星承制方提供所需的操作窗口和透波窗口。卫星承制方要求的特定窗口的大小和位置由双方协调确定。
5.2 电气接口
5.2.1 分离电连接器
应明确星箭间的分离电连接器的下列特性：
a.连接器的型号；
b.用户可用的芯数(接点数)；
c.卫星电源与火工装置点火电路之间的隔离；
d.电气性能；
e.屏蔽要求；
f.锁定标志。
5.2.2 脐带链路
星箭对接后，卫星与地面电气支持设备(EGSE)之间的脐带链路特性如下：
a.链路的数量与形式。
b.限值参数：
最大电压(V)；
最大电流(A)；
功率(W)；
单路最大电阻或压降(Ω或V)。
c.工作约束条件：
功能数量；
功能类型。
d.连接器接点电流。
e.一致性确认：
端—端电阻(Ω)；
线—地绝缘性；
线—线绝缘性。
5.2.3 卫星专用电气指令
卫星专用电气指令包括标准指令和备用指令，由运载火箭产生，供卫星专用。应列出指令清单，并对其相关特性作出说明。
5.2.3.1 火工装置指令
应对与火工装置指令相关的电路作一般说明，并给出对应电路示意图以及下列主要特性：
a.指令数；
b.指令电压(V)；
c.指令脉宽；
d.输出绝缘性(Ω)；
e.火工装置负载电流(A)；
f.两个指令间的间隔；
g.导线与结构间的绝缘性；
h.对卫星设备的保护；
i.卫星上的使用限制。
5.2.3.2 监视演示回路指令
应对与监视回路指令相关的电路作一般说明，并给出对应电路示意图以及下列主要电特性：
a.在地面或飞行中的可用指令数；
b.指令电压(V)；
c.指令持续时间；
d.电阻(Ω)；
e.星上电路绝缘性；
f.卫星上的保护措施；
g.卫星上的使用限制(最大电压和电流)。
5.2.3.3 其他电气指令
应对与这些指令相关的电路作一般说明，并给出对应的电路示意图以及下列主要特性；
a.地面或飞行中的可用指令数；
b.输出电压(V)及火工品负载电流(A)；
c.卫星上的保护措施；
d.电路与结构之间的绝缘性；
e.电磁兼容性(EMC)应符合GJB 151A的规定；
f.卫星上的使用限制。
5.2.4 分离状态传输
分离状态信号由运载火箭遥测系统传输，可用于分离状态传输的方式主要有：
a.微动开关；
b.断线；
c.监视回路；
d.其他。
5.2.5 飞行中的遥测
运载火箭承制方应对运载火箭飞行遥测获得的卫星数据类型进行说明，主要有：
a.星箭接口处的力学和热环境数据；
b.卫星要求的特定内部测量参数。
5.2.6 电源
应给出与电源相关电路的一般说明，并给出对应电路示意图、电源主要电特性及其允差：
a.电压(V)；
b.电流(A)；
c.卫星上的绝缘；
d.卫星上的保护措施；
e.EMC。
5.2.7 地电位连续性
应对卫星相对于地电位的电连续性要求进行说明：
a.卫星参考点的位置(可为多点)；
b.卫星金属件与最近参考点之间的最大电阻；
c.星箭分离面最大电阻。
5.3 射频/电磁接口
对卫星、运载火箭和发射场的辐射应一起进行评估，以确定兼容性和可能的限制条件。5.6条和5.7条规定了验证这些接口的方法。
5.3.1 RF遥测(TM)和指令链路
运载火箭承制方应通过电连接器为卫星TM和指令天线与卫星系统测试设备提供TM和指令链路。
该链路的建立和使用将是通过与卫星整流罩有关的RF窗口(即透波窗口)、接收天线或其他等效方法而实现。除了因运载火箭的工作而使RF传输受到限制以外，从卫星在运载火箭上安装封闭工作结束直到发射，该链路都应是可用的。卫星的TM和指令频率必须与运载火箭的频率不相重迭。
运载火箭承制方应说明运载火箭的以下特性：
a.遥测(TM)：
带宽；
频率；
发射机功率。
b.自毁指令接收机(CDR)：
带宽；
频率。
c.雷达应答机(RT)：
带宽；
频率；
峰值功率。
d.RF窗口的特性：
材料；
部位(包括方位、位置和尺寸)；
频率范围插入损失。
5.3.2 星箭电磁兼容性
a.卫星应与运载火箭产生的电磁场兼容，运载火箭的电磁环境见5.5.6.1条；
b.运载火箭应与卫星产生的电磁场兼容，卫星的电磁环境见5.5.6.2条。
5.4 任务性能
本条规定了运载火箭性能参数，以确保运载火箭把卫星送入预定的轨道。
5.4.1 一般轨道
5.4.1.1 性能
运载火箭发射一般轨道卫星的性能可以用运载能力与轨道参数(hP、hn、i、ω)的图形关系表示。发射一般轨道卫星的运载能力的实例见附录C(参考件)中的图C1～C3。
5.4.1.2 运载火箭入轨精度
应给出一般轨道参数的标准偏差(1σ)或最大偏差(3σ)：
a.轨道倾角偏差Δi(°)；
b.近地点高度偏差ΔhP(km)；
c.远地点高度偏差Δha(km)；
d.近地点幅角偏差Δω(°)；
e.升交点赤经偏差ΔΩ(°)。
卫星入轨精度(3σ)的实例见附录C中表C1。
5.4.1.3 发射窗口
当运载火箭和发射场对发射窗口存在约束条件时，则应予以明确。
5.4.2 地球同步转移轨道(GTO)。
5.4.2.1 性能
5.4.2.1.1 应根据运载火箭把卫星送入轨道的参数确定地球同步转移轨道(GTO)：
a.轨道倾角i(°)；
b.近地点高度hP(km)；
c.远地点高度hn(km)；
d.近地点幅角ω(°)；
e.升交点赤经Ω(°)。
注：半长轴a和轨道偏心率e可从几个轨道要素的相互关系中导出。
5.4.2.1.2 运载火箭在地球同步转移轨道的性能可以用下述两种情况表示：
a.采用标准的PLA，运载能力为单星、双星或多星发射的质量。
b.采用非标准的PLA，则运载能力为卫量和PLA的质量。
5.4.2.1.3 对于发射地球同步转移轨道(i＝28.5°)卫星，远地点高度偏差对运载能力的影响的实例见附录C中图C4。发射地球同步轨道卫星运载能力与轨道倾角的关系的实例见附录C中图C5。应给出亚同步和超同步地球同步转移轨道详细的性能计算。
5.4.2.2 运载火箭入轨精度
应根据地球同步转移轨道(GTO)参数来说明运载火箭把卫星送入地球同步转移轨道(GTO)的入轨精度。
给出轨道参数的标准偏差(1σ)或最大偏差(3σ)。提供相关参数的协方差矩阵或一组等效的运载火箭状态参数：
a.轨道倾角偏差Δi(°)；
b.近地点高度偏差ΔhP(km)；
c.远地点高度偏差Δha(km)；
d.近地点幅角偏差Δω(°)；
e.升交点赤经偏差ΔΩ(°)。
5.4.2.3 发射窗口
当运载火箭或发射场对发射窗口存在约束条件时，则应表示为标准发射窗口的一部分。
5.4.3 太阳同步轨道(SSO)
5.4.3.1 性能
发射太阳同步轨道卫星的运载火箭性能用运载能力与轨道高度的图形关系表示，其实例见附录C中图C6。
5.4.3.2 运载火箭的入轨精度
应给出太阳同步轨道参数的标准偏差(1σ)或最大偏差(3σ)：
a.近地点高度偏差ΔhP(km)；
b.偏心率偏差Δe；
c.轨道倾角偏差Δi(°)；
d.轨道周期偏差ΔT(s)；
e.近地点幅角偏差Δω(°)。
5.4.3.3 发射窗口
按照卫星轨道的当地时间和运载火箭的性能给出太阳同步轨道发射窗口的约束条件。
5.4.4 卫星定向和分离
5.4.4.1 一般描述
运载火箭把卫星送入轨道后对姿态机动能力的一般说明应包括：
a.姿态控制系统类型和推进剂(气体、肼)；
b.针对各类卫星的不同要求(三轴稳定、自旋稳定、多次分离)，运载火箭从入轨到飞行任务结束过程的典型时序；
c.运载火箭的约束条件。
5.4.4.2 定向性能
规定轨道坐标系，以便确定自旋稳定卫星的纵轴定向或三轴稳定卫星的二轴定向的方向要求：
a.当运载火箭存在这种能力时，为卫星方向随时间而变化(作为发射时间的函数)提供必要的数据；
b.自旋性能包括最大自旋速度及其误差；
c.三轴稳定卫星的指向精度(卫星分离前三轴的姿态误差)和自旋稳定卫星的指向精度(卫星分离后的动量矢量的指向误差)；
d.运载火箭分离系统提供的最小相对分离速度。
星箭分离结束时，卫星的初始姿态角偏差值和初始姿态角速度偏差值(3σ)的实例见附录C中的表C2。
5.5 诱导环境和极限载荷
5.5.1 极限载荷
运载火箭承制方应给出以下极限载荷。
5.5.1.1 稳态载荷
运载火箭飞行过程中稳态载荷以卫星质心或星箭对接面处的加速度表示，其实例见附录D(参考件)中图D1。
5.5.1.2 准静态载荷
运载火箭飞行过程中准静态载荷为稳态载荷与动态载荷之代数和，以卫星质心或星箭对接面处的加速度表示，其实例见附录D中的D2章。
5.5.1.3 最大极限载荷
规定PLA能承受的最大允许载荷。
5.5.2 力学环境
5.5.2.1 低频振动
运载火箭承制方应根据有关频段的正弦振动和瞬态振动在星箭对接面上的响应，给出等效正弦振动谱型，其实例见附录D中图D2～D3。
5.5.2.2 随机振动
运载火箭承制方根据随机振动载荷的功率谱密度(g2/Hz)，给出随机振动三轴方向飞行的包络谱，其实例见附录D中图D4。
5.5.2.3 噪声
运载火箭承制方给出卫星整流罩内或支承结构的飞行噪声谱，其实例见附录D中图D5。并表明卫星与卫星整流罩的填充因子(卫星与整流罩容积之比)。
5.5.2.4 冲击
规定星箭接口处及有关区域的最高冲击谱，其实例见附录D中图D6。
5.5.2.5 卫星质心位置限定
标明卫星质心的位置，说明卫星质心位置与PLA重量的限定，其实例见附录D中第D3章。
5.5.3 热环境
5.5.3.1 一般要求
卫星热环境包括以下方面：
a.卫星从总装测试厂房至发射场发射区转运阶段的热环境；
b.星箭对接至发射前阶段的热环境；
c.卫星发射至星箭分离阶段的热环境。
5.5.3.2 地面操作热环境
地面操作热环境主要有：
a.工作环境温度；
b.相对湿度；
c.整流罩内气体流速(m/s)；
5.5.3.3 运载火箭飞行时的热流
运载火箭承制方应给出运载火箭飞行时整流罩内表面典型参考点的热流随时间变化曲线，其实例见附录D中D4章；对于返回式卫星应给出运载火箭飞行时卫星外表面热流随时间变化曲线。
5.5.3.4 抛卫星整流罩时刻的热流
运载火箭承制方应给出抛卫星整流罩时，自由分子流加热的最大热流值。
5.5.3.5 星箭分离阶段的热流
运载火箭承制方给出星箭分离阶段运载火箭在卫星上产生的最大热流及持续时间。
5.5.4 卫星整流罩内静压力
运载火箭承制方给出运载火箭飞行时，卫星整流罩内静压力随时间变化曲线，其实例见附录D中图D7。
5.5.5 污染和洁净度
5.5.5.1 卫星的污染
运载火箭承制方应给出在下述卫星操作过程中，运载火箭材料出气和分离系统对卫星产生的有机及微粒沉积物：
a.卫星在整流罩(或支承结构)内的地面阶段(转运、射前)和飞行阶段；
b.运载火箭上产生的烟羽流。
5.5.5.2 洁净度
运载火箭承制方应给出与卫星有关的操作间、转运容器和卫星整流罩内空气的洁净度等级。
5.5.6 无线电和电磁环境
5.5.6.1 运载火箭产生的环境
运载火箭承制方应说明附加辐射干扰量级：
a.运载火箭窄带电场发射的附加辐射，其实例见附录D中图D8；
b.运载火箭宽带电场发射的附加辐射，其实例见附录D中图D9；
c.运载火箭窄带磁场发射的附加辐射，其实例见附录D中图D10。
5.5.6.2 卫星产生的环境
运载火箭承制方应说明其在窄带电场可接受的附加辐射量级。
5.6 验证分析和文件
本条规定验证与5.1～5.5条有关的星箭特定接口的过程。
5.6.1 验证定义
验证分析包括下列任务：
a.计算相应型号的运载火箭把卫星送入目标轨道的参数，确定相关的发射窗口。
b.分析星箭分离后的姿态指向以及星箭的分离间距。
c.进行卫星验证试验，获取必要的数据，并整理试验结果，包括力学环境、热环境、无线电兼容性和污染分析等。
d.验证星箭在联合操作和飞行期间的安全性。
e.利用运载火箭遥测数据，应对下述典型的飞行事件进行发射轨道和飞行环境的评定分析：
发动机点火；
推力中止；
级间分离；
抛卫星整流罩；
卫星分离之前的运载火箭稳定段；
星箭分离。
5.6.2 验证分析方法
验证分析方法如下：
a.数学模型和仿真；
b.概率计算和统计估计；
c.技术系统的性能评价；
d.分析和整理研究结果；
e.计算结果与已知类似数据的比较。
5.6.3 验证分析阶段
5.6.3.1 可行性分析
估计运载火箭和卫星在特定技术方面之间基本兼容的条件，可行性分析由卫星和运载火箭承制方之间共同进行。
5.6.3.2 初步分析
运载火箭和卫星承制方在进行各个验证试验之前，应提供参考输入数据并进行数学模型分析。
5.6.3.3 最终分析
以运载火箭和卫星承制方所提供的最后发射的技术状态和试验验证的数学模型为基础，用输入的数据完成最终分析。
5.6.4 分析使用的输入数据
卫星承制方应向运载火箭承制方提供卫星全部技术状态的下列输入数据：
a.预定的轨道、分离姿态和自旋速度(有应用时)；
b.质量特性；
c.理论外形图；
d.模态特性的动力学模型；
e.热模型；
f.规定的机械接口和电气接口；
g.无线电频率特性和传输图；
h.在地面和运载火箭/卫星联合操作期间的电磁辐射特性。
卫星承制方和运载火箭承制方在坐标系一致的情况下，提供数据资料和数学模型。数据资料用标称值和有关允差说明。
卫星承制方和运载火箭承制方的协议书中应规定进行验证分析所要求的任何附加内容。
5.6.5 验证分析
为了确保卫星与运载火箭环境的兼容性，运载火箭承制方应采用卫星承制方所提供的输入数据完成下述分析。
5.6.5.1 弹道和性能分析
运载火箭承制方应计算运载火箭发射轨道及卫星要求轨道的有关性能及其余量。分析应包括下列项目：
a.飞行程序和飞行轨迹的说明；
b.运载火箭主要参数表；
c.性能余量和相关的概率；
d.运载火箭跟踪站的可跟踪性；
e.卫星任务约束条件的验证。
5.6.5.2 发射窗口分析
当发射窗口的约束条件与运载火箭性能或运载火箭系统有关时，运载火箭承制方应对卫星每天可利用的发射时间(发射窗口)进行计算。最终得到的发射窗口应与卫星提供的发射窗口相吻合。
5.6.5.3 定向、分离分析
运载火箭承制方应对卫星定向要求的姿态进行分析，必要时，还应对预定的卫星自旋速度进行分析。应根据星箭分离后卫星运动学的情况，进行卫星的分离过程和动力学的描述。
5.6.5.4 载荷耦合分析
运载火箭承制方进行载荷耦合分析，预测卫星的载荷和星箭之间的相对位移。卫星承制方使用分析结果验证卫星设计与运载火箭环境之间的兼容性，调整卫星的振动试验量值。
对所有关键性的地面和飞行工况，应计算耦合的星箭结构纵向和横向载荷。运载火箭承制方应给出卫星模型选定结点处的力、加速度、相对位移和时间历程，以及星箭接口处的力、加速度和时间历程。卫星承制方应按一致同意的分析格式，计算卫星内部响应。
5.6.5.5 间隙分析
运载火箭承制方应分析研究在运载火箭飞行和卫星分离期间，卫星与整流罩或其它结构之间的关键间隙。
5.6.5.6 热分析
为了满足卫星系统允许的热约束条件，运载火箭承制方进行热环境分析，预测卫星在地面工作和飞行期间的温度。卫星承制方可以利用分析结果验证卫星分系统与运载火箭环境的一致性。
应预测从卫星装入整流罩内一直到星箭分离时，运载火箭对卫星的热环境影响。应计算整流罩分离时的气动热流。
5.6.5.7 射频(RF)链路分析
完成卫星和发射场区各种地面设施之间的射频(RF)链路分析，以便确定射频(RF)链路的余量。
5.6.5.8 电磁兼容性和干扰分析
运载火箭承制方应进行星箭在发射场区相关设施之间的电磁兼容性(EMC)及电磁干扰分析。必要时，应进行验证试验。
运载火箭承制方应向卫星承制方提供在星箭联合地面操作期间、飞行期间及星箭分离前、后在整流罩内卫星附近的电磁辐射特性。
5.6.5.9 污染分析
完成运载火箭和卫星之间在各分离阶段和分离后的污染分析。
5.6.6 安全性
运载火箭承制方利用卫星提供的数据(包括卫星高压容器和有关部件的安全系数及强度余量)进行地面联合操作和发射期间的安全性验证。
地面联合操作期间应考虑下列两种情况：
a.卫星周围有人或危险物品；
b.卫星周围无人或危险物品。
5.6.7 发射准备状态评估
为了实施发射工作，由运载火箭承制方和卫星承制方共同验证分析结果，对有关试验作出评估。在星箭接口控制文件中规定相应的程序。
5.6.8 发射后评价
发射后，运载火箭承制方应分析运载火箭遥测数据，提出飞行评价报告：
a.发射前准备和发射操作以及运载火箭飞行时序的执行情况(包括抛整流罩、姿态定向和分离等)；
b.轨道参数评价；
c.发射前操作和飞行期间的诱导环境，并说明异常现象。
5.6.9 文件
5.6.9.1 星箭接口要求文件
卫星承制方应提出接口要求文件，其主要内容有：
a.卫星的技术特征；
b.卫星的轨道及其偏差；
c.卫星的分离姿态、自旋速度及其偏差；
d.卫星的约束条件；
e.卫星的发射窗口；
f.要求的轨道服务(有应用时)。
5.6.9.2 星箭兼容性的验证计划
运载火箭和卫星管理部门共同准备文件，内容见附录E(参考件)中表E1，主要有：
a.验证项目清单；
b.验证剪裁的原则和策略；
c.验证工作程序；
d.验证工作进度。
5.6.9.3 接口控制文件
运载火箭承制方在接口要求文件的基础上，准备接口控制文件。文件包括确定和验证星箭之间所需要的兼容性全部技术要求，最终由卫星承制方确认。