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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
