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Small Spacecraft Mission Service User’s Manual – Proof of Concept flight on

VEGA DC/BD/ST/CDU N17-003

ESA & Arianespace Proprietary Information

Page : i

SMALL SPACECRAFT MISSION SERVICE

USER MANUAL

for

SSMS PROOF OF CONCEPT FLIGHT ON VEGA

Issue 1 Revision 0 / Feb. 2017

The signatures below certify the validity of this document

ARIANESPACE ESA

Prepared by: Claude DUPUIS

Prepared by: Aldo SCACCIA

Approved by:

Approved by : Fabio CARAMELLI



RECORD OF AMENDMENTS

Issue - Revision

Date Modified Pages AUTHORIZATION

AE ESA 1.0 Feb. 2017 All



ESA & Arianespace Proprietary Information Page : 3 of 26

TABLE OF CONTENTS

1. Introduction …………………………………………………………………………………… 4

1.1. Purpose of the present User Manual 1.2. Definition, Acronyms and Abbreviations

2. Technical description…………………………………………………………………..…… 6

2.1. SSMS dispenser description 2.2. Small Spacecraft accommodation 2.3. Reference configurations for the SSMS PoC flight

3. Mission requirements ……………………………………………………………………… 11

3.1. Orbit 3.2. Launch time 3.3. Launch sequence, launch duration 3.4. Conditions at S/C separation 3.5. Small Satellite sequence after separation

4. Interface requirements ……………………………………………………………………13

4.1. Interface requirements for Mini satellite (Class#1) 4.2. Interface requirements for Micro satellite (Class#2) 4.3. Interface requirements for Nano satellite (Class#3) 4.4. Interface requirements for Cubesats deployers (Class#4) 4.5. Interface verifications

5. Design requirements, Flight Environment & S/C Qualification …………… 21 5.1. Design requirements for Mini satellite (Class#1) 5.2. Design requirements for Micro satellite (Class#2) 5.3. Design requirements for Nano satellite (Class#3) and Cubesats deployer

(Class#4) 5.4. Flight Environment 5.5. S/C qualification

6. Stack integration & Launch campaign ……………………………………………… 24

6.1. Stack integration flow 6.2. S/C checks/accessibility after Stack integration 6.3. S/C hazardous propellant fuelling after Stack integration




1. Introduction

1.1. Purpose of the present User Manual The present User’s Manual for Small Spacecraft is dedicated to the SSMS PoC flight on VEGA. It is intended to provide basic information for Smalls Passengers on this flight using the SSMS hardware and processes on a Vega launch system operated from the Guiana Space Centre. The content encompasses:

the classification of Small Spacecraft; the SSMS dispenser description; the envisaged mission; the description of the interfaces between Small Spacecraft and Launch Vehicle (LV); the requirements for Small Spacecraft design and verification; the ground operations for stack integration and launch.

Together with the Vega User’s Manual, it gives readers the information to assess the compatibility with the proposed standardized SSMS configurations.




1.2. Definitions, Acronyms and Abbreviations CLA Coupled Loads Analysis CoG Centre of Gravity COTE Check Out Terminal Equipment DUA Demande d’Utilisation Arianespace (Application to use Arianespace Launch vehicles) ESA European Space Agency GSE Ground Support Equipment ICD Interface Control Document LLL Low cost Launch opportunities for Light missions LMP-103S “green” Liquid MonoPropellant LSA Launch Services Agreement LTAN Local Time of Ascending Node LV Launch Vehicle MLB Motorized Light Band MUV Manuel Utilisateur Vega (Vega User’s Manual) N2H4 Hydrazine N/A Non Applicable PoC Proof of Concept PSC Planetary Systems Corporation QSL Quasi Static Loads RfW Request for Waiver S/C Spacecraft SSMS Small Spacecraft Mission Service SSO Sun Synchronous Orbit Stack Small spacecraft + adapters + SSMS dispenser TBC To Be Confirmed TBD To Be Defined UCIF Upper Composite Integration Facility VESPA Vega Secondary Payload Adapter




2. Technical description 2.1. SSMS Dispenser description

The SSMS dispenser has a modular design allowing to aggregate different classes of small spacecraft. It is composed of several modular elements, which are briefly described here below: Tower module

Column Module

Lower Module

The lower module is the common core of the SSMS dispenser. It has a hexagonal part and a platform. Six positions are available on the hexagonal part to accommodate up to 6 very small spacecraft or up to 12 (12U) Cubesats deployers horizontally (“lower module positions”). On the platform, up to four positions are available to accommodate small spacecraft (“Deck 2 position”). Then, two, three of four tower modules can be installed on top of the platform to accommodate small spacecraft (“Deck 1 position”). Similarly, a central column module can be installed to accommodate another bigger small spacecraft (“Top position”). Several heights are available for the central column. The lower module is itself attached to the launcher using the existing 937 conical structure. Note: The SSMS dispenser is in its development phase and the definition of these elements may still evolve.




With these elements, many different Stack configurations can be envisaged:

Piggyback configuration

One (1) Main passenger

Six (6) Nano satellites or Six (6) Cubesats deployers

Constellation configuration

Four (4) identical Micro satellites


“FLEXI-3” configuration

One (1) Mini satellite

Three (3) Micro satellites


“FLEXI-4” configuration

One (1) Mini satellite

Height (8) Micro satellites





2.2. Small Spacecraft accommodation The available accommodation and Launch Service depend on the mass and volume of the Small Spacecraft Passenger. Accommodation dimensions as a function of standard position has been established in relation to S/C classes (see AO).

# Class Mass range Standard Envelope Volume Standard position (Kg) (mm)

1 Mini Satellite 400 – 200 H1800 x L1200 x W1200 Top position

2 Micro Satellite 200 – 60

H1200 x L800 x W800

Top position or

Deck#2 position (FLEXI-3 configuration)

H1100 x L700 x W600

Deck #1 position (FLEXI-4 configuration)

or Deck#2 position (FLEXI-4

configuration)

3 Nano Satellite 60 – 25 H600 x L400 x W300

Deck#1 position (FLEXI-4 configuration)

or Base module position

4 Cubesats Deployer 25 – 1 H530 x L320 x W280

(doors open) Base module position




2.3. Reference configurations for the SSMS PoC flight For the SSMS PoC flight on Vega, the envisaged stack configurations correspond to the so-called FLEXI-3 or FLEXI-4 configurations, allowing to embark up to fifteen (15) Small Spacecraft or Cubesats deployers. The sketches below illustrate the envisaged Stack configurations:

SSMS Stack conf#1 SSMS Stack conf#2

FLEXI 4 configuration with Mini and Micro S/C

FLEXI 4 configuration with Micro S/C

One (1) Mini satellite on a stretched central column Four (4) Micro satellites on Deck#1 on four tower modules Four (4) Micro satellites on Deck#2 Three (3) Nano satellites on Base module Three (3) Cubesats deployers on Base module

One (1) Micro satellite on the central column Four (4) Micro satellites on Deck#1 on four tower modules Four (4) Micro satellites on Deck#2 Three (3) Nano satellites on Base module Three (3) Cubesats deployers on Base module




SSMS Stack conf#3 SSMS Stack conf#4

FLEXI 4 configuration with Mini, Micro and Nano S/C

FLEXI 3 configuration with Micro S/C

One (1) Mini satellite on a stretched central column Four (4) Nano satellites on Deck#1 on four tower modules Four (4) Micro satellites on Deck#2 Three (3) Nano satellites on Base module Three (3) Cubesats deployers on Base module

One (1) Micro satellite on the central column Three (3) Micro satellites on Deck#2 Three (3) Nano satellites on Base module Three (3) Cubesats deployers on Base module

As the modular design of the SSMS dispenser allows a large number of possible combinations, alternative configurations can be also envisaged depending of customer demands. The actual configuration of the SSMS PoC flight on Vega will be frozen following extensive analysis of customer demands.




3. Mission requirements 3.1. Orbit The reference orbit for the SSMS PoC flight mission is Sun Synchronous Orbit (SSO) in the 500 – 800 km altitude range. Other orbit might be considered within the Vega launch system capability, i.e. any Low Earth Orbit (LEO) in the 400 – 1400 km altitude range. Thanks to the reignition capability of the VEGA last stage (AVUM), the small passengers can be separated on up to two different altitudes (with the corresponding inclination for a SSO). The Local Time of Ascending Node (LTAN) can be either 10:30 a.m. or 06:00 a.m. The actual targeted orbit and Local Time of Ascending Node for the SSMS PoC flight will be frozen following SC aggregate definition. 3.2. Launch time For SSO mission, the launch window is reduced to a single launch time. This launch time will be determined taking into account the targeted Local Time of Ascending Node (LTAN). 3.3. Launch sequence, launch duration The overall launch sequence and the separation time of the Small Spacecraft will be determined depending of the passenger mission requirements, of the available ground stations, etc … A typical ascent and separation sequence is the following:

Three-stage ascent, AVUM boost(s) to reach the targeted orbit (orbit #1) of the top position passenger or

Deck#1 passengers, When present, longitudinal separation of the S/C on top position, Avoidance maneuvers, AVUM boosts (as necessary) to reach the targeted orbit (orbit #2) of the Deck#1

and/or Deck#2 passengers, Longitudinal separation of the S/C on Deck#1. It can be a simultaneous separation of

all the S/C or a separation two by two, Avoidance maneuvers (as necessary), Longitudinal separation of the S/C on Deck#2. Similarly to Deck#1, it can be a

simultaneous separation of all the S/C at the same time or a separation two by two. Additionally, when necessary, a tilt can be envisaged (S/C separation axis up to 10° from launcher longitudinal axis),

Avoidance maneuvers (as necessary), Lateral separation of the S/C on Lower module. It can be a simultaneous separation

of all the S/C at the same time or a separation two by two, or a separation three by three,

Avoidance maneuvers,




Final AVUM boost (as necessary) to trigger a controlled re-entry of the remaining parts of the AVUM stage into the Earth atmosphere.

The maximum delay between lift-off and a spacecraft separation can be up to 7 hours.

3.4. Conditions at S/C separation The separation conditions will be determined in the frame of the mission analysis, based on spacecraft mass properties (nominal and dispersed values). S/C dispersions on CoG positions shall be reduced to its minimum possible extent. Actual S/C mass properties shall be measured. 3.5. Small Satellite sequence after separation The Small Satellite sequence of events after separation from the launch Vehicle shall respect the following timeline requirements:

- To ensure safe distance between the bodies, small S/C will be asked to implement a delay for the activation of the S/C attitude control and/or any deployment. This delay will be determined in the frame of the mission analysis, by the separation analysis.

- To ensure compatibility with AVUM avionics, small S/C will be asked to implement a delay for switching-on any TM transmitter. This delay will be determined in the frame of the mission analysis, by the EMC analysis.




4. Interface requirements This chapter defined the requirements in terms of usable volume, mechanical interface, electrical interface and interfaces verification. For the SSMS PoC flight on Vega, the S/C-LV interface is based on Lightband system from PSC. Other adapter can be envisaged. 4.1. Interface requirements for Mini Satellite (Class#1) S/C Position Requirement Reference Values Negotiable? Notes

TOP

Usable Static Volume

H 1800 mm 1200 x 1200 mm

yes

Usable volume can be extended up to:

Separation Direction Longitudinal no

S/C hazardous propellant fuelling

Allowed n/a

S/C Mechanical Interfaces Requirement Reference

Values Negotiable? Notes

PSC Mark II Lightband

24-inches

S/C Interface Ring External Diameter 665 mm no

S/C Interface Ring Internal Diameter 589 mm no

S/C Interface Ring mass 1.098 kg ±5% no

Interface screws holes diameter

609.6 mm ± 0.25 mm no

Fasteners type Allen Head cap screws M6 yes

The fastener dimension can be reduced

Tightening torque 11 - 13 Nm yes Minimum admissible torque: 8 Nm

Adjoining surface planarity 0.16 mm yes

Depending on the stiffness of the adjoining structure. The adjoining structure stiffness shall guarantee the planarity in all loads conditions.

PSC Transition Ring (Optional)

24-inches

S/C Transition Ring External Diameter 665 mm TBC no

Optional for flight S/C Transition Ring Internal Diameter 589 mm TBC no

S/C Transition Ring mass TBD no

Interface screws 609.6 mm ± no




holes diameter 0.25 mm




S/C Electrical interfaces Requirement Reference


PSC

Connectors

Number Up to 6 no

Size 15 pins no

Lines available before Stack

integration on LV

Number of Available lines for Class#1 S/C

≤ 74 no

Lines available after Stack

integration on LV (including

countdown phase on launch

pad)


≤ 34 yes Depending on Stack configuration




4.2. Interface requirements for Micro Satellite (Class#2) S/C Position Requirement Reference Values Negotiable? Notes

TOP


H 1200 mm 800 x 800 mm

yes

For TOP position, the usable volume can be extended up to:

Deck#2 Separation Direction Longitudinal no


Allowed n/a

The S/C fuelling ports (if needed) accessible after Stack integration i.e. shall be located on one of the S/C faces facing the fairing.

Deck#1


H 1100 mm 700 x 600 mm

Deck#1: yes Deck#2: no

Protrusions out of the usable volume can be negotiated on a case by case basis.


For Deck#2 position, a tilt can be envisaged (S/C separation axis up to 10° from launcher longitudinal axis).

Deck#2


Allowed n/a

The S/C fuelling ports (if needed) accessible after Stack integration i.e. shall be located on the S/C face facing the fairing.







13-inches





330.2 mm ± 0.25 mm no






PSC Transition Ring (Optional)

13-inches

S/C Transition Ring External Diameter 382 mm TBC no

Optional for flight S/C Transition Ring Internal Diameter 308 mm TBC no

S/C Transition Ring mass TBD no


330.2 mm ± 0.25 mm no






PSC

Connectors

Number Up to 4 no

Size 15 pins no


integration on LV


≤ 60 no




pad)


≤ 12 yes Depending on Stack configuration




4.3. Interface requirements for Nano Satellite (Class#3) S/C Position Requirement Reference Values Negotiable? Notes

Deck#1


H 600 mm 400 x 300 mm yes



S/C fuelling Not Allowed yes

Depending on Stack integration flow. Note: Anyhow, loading of not-hazardous propellant (Xenon for instance) is allowed prior to S/C integration on SSMS dispenser.

LOWER

Module Usable Static Volume

H 600 mm 400 x 300 mm yes


Separation Direction Lateral no


Not Allowed no

Note: Anyhow, loading of not-hazardous propellant (Xenon for instance) is allowed prior to S/C integration on SSMS dispenser.


Values Negotiable Notes


8-inches





203.2 mm ± 0.25 mm no

Fasteners type Allen Head cap screws M6 (TBC)

yes The fastener dimension can be reduced.







PSC Transition Ring

8-inches

S/C Interface Ring External Diameter 255 mm TBC no

Optional for flight S/C Interface Ring Internal Diameter 178 mm TBC no

S/C Interface Ring mass TBD no


203.2 mm ± 0.25 mm no

Fasteners type Allen Head cap screws M6 (TBC)

yes The fastener dimension is suggested

Tightening torque 11 - 13 Nm yes



PSC

Connectors

Number Up to 2 no

Size 15 pins no


integration on LV


≤ 30 no




pad)


≤ 12 no




4.4. Interface requirements for Cubesats deployers (Class#4) S/C Position Requirement Reference Values Negotiable? Notes

LOWER

Module


H 530 mm 320 x 280 mm no

Separation Direction Lateral no


Not Allowed no



Bolted interface

With inserts on SSMS lower

module

Number of bolts 12 Yes

Depending of cubesats deployer manufacturer and size

Type of bolts M6 Pitch 1.0 mm

Insert length Min: 9.0 mm Max: 12.0 mm




integration on LV Number of available lines None no


integration on LV Number of available lines None no




4.5. Interfaces verification Dimensional fit-check: Volume compatibility between Small Spacecraft and SSMS dispenser. Nominally the volume compatibility check between Small spacecraft and the SSMS dispenser will be done based on CAD models. S/C CAD models shall represent the as-built S/C configuration. It shall include provisions for any MLI, harness routing, etc… It shall also include non-flight items when these elements are to be removed after S/C integration on the SSMS dispenser. As an option (to be ordered by the customer), especially when the S/C have protrusions outside the defined usable volume, a volume fit-check may be organized prior to the Stack integration and launch campaign. In such case, the Structural Model of the SSMS dispenser will be used, and the customer shall make available S/C representative dummy. Mechanical and electrical fit-check: Interface compatibility between Small Spacecraft and SSMS dispenser. Nominally, the mechanical compatibility check will be done using flight standard metal (Al alloy) plates (S/C Transition Ring - see above section 4.1, 4.2 & 4.3), mounted on both sides, to guarantee compliance on both sides with required mechanical tolerances. Note: A separation test is part of the standard acceptance test of each Lightband hardware.




5. Design requirements, Flight Environment & S/C Qualification This chapter defined the design requirements, the flight environment and the S/C qualification logic. 5.1. Design requirements for Mini satellites (Class#1) S/C Design Requirements

Requirement Reference Values Negotiable? Notes

Masses, CoG and Inertia

CoG Position: distance from separation plane

≤ 500 mm yes Subject to Stack configuration

Static unbalance: distance “d” between CoG and the longitudinal axis

d ≤ 10 mm yes

Roll Inertia ≤ 65.0 kgm2 yes Pitch and Yaw Inertia ≤ 130.0 kgm2 yes

Propellant Sloshing Sloshing mass vs. total fluid mass shall not exceed 5%

no

Main Frequencies Lateral frequencies > 30 Hz yes

Longitudinal frequencies > 60 Hz yes

QSL at S/C CoG

Longitudinal* - 8.5 g + 5 g

yes

Lateral** ± 3 g yes Depending on Stack configuration and relevant CLA

*: The minus signs indicate compression along the longitudinal axis and the plus signs tension. The gravity load is included **: The lateral load on payloads may act in any direction in the range [0-360°]. Moreover they must be considered simultaneous and uncorrelated 5.2. Design requirements for Micro satellites (Class#2) S/C Design Requirements

Requirement Reference Values Negotiable? Notes





d ≤ 10 mm no

Propellant Sloshing Sloshing mass vs. total fluid mass shall not exceed 5%

no

Main Frequencies Lateral frequencies > 60 Hz yes

Longitudinal frequencies > 90 Hz yes

QSL at S/C CoG

Longitudinal* - 8.5 g + 5 g

yes

Lateral** ± 5 g yes Depending on Stack configuration and relevant CLA

*: The minus signs indicate compression along the longitudinal axis and the plus signs tension. The gravity load is included **: The lateral load on payloads may act in any direction in the range [0-360°]. Moreover they must be considered simultaneous and uncorrelated




5.3. Design requirements for Nano satellites (Class#3) and Cubesat deployers (Class#4) S/C Design Requirements

Requirement Reference Values

Negotiable? Notes





d ≤ 10 mm no

Main Frequencies Lateral frequencies > 115 Hz yes Down to 90 Hz Longitudinal frequencies > 115 Hz yes Down to 90 Hz

QSL at S/C CoG Longitudinal* - 14.5 g + 10.5 g yes

Depending on Stack configuration and relevant CLA Lateral** ± 3 g

Longitudinal and Lateral axes of the Launcher are referred: For S/C installed on the lower module, this implies that the longitudinal loads are applied in lateral direction for the spacecraft. *: The minus signs indicate compression along the longitudinal axis and the plus signs tension. The gravity load is included **: The lateral load on payloads may act in any direction in the range [0-360°]. Moreover they must be considered simultaneous and uncorrelated

1.1. 5.4. Flight Environment Requirement Reference Values Negotiable? Notes Sine Environment

Sine vibrations

Lateral direction (Flight limit levels)

5 – 70 Hz : 2.0 g 70 – 110 Hz : 1.0 g 110 – 125 Hz : 0.2 g

yes

Depending on Stack configuration and relevant CLA

Longitudinal direction (Flight limit levels)

5 – 70 Hz : 2.0 g 70 – 110 Hz : 1.0 g 110 – 125 Hz : 0.2 g

Qualification factor 1.25 no

Random Environment

Random vibrations

Flight limit levels 20 to 100 Hz: 0.025 g2/Hz 100 to 2000 Hz: –3 dB/octave

Yes Qualification factor

2.25 (factor by which to multiply the Power Spectral Density)

Duration 2 min Acoustic Environment

Acoustic vibrations

OASPL Overall Acoustic Sound Pressure Level

Refer to Vega User’s Manual no

Shock Environment

SRS (Shock response spectrum)

For LV events Covered by S/C separation n/a For S/C separation event

100-1000 Hz 1000-10000 Hz yes 40-1000 g 1000-800g

Static pressure in the fairing Static pressure in the fairing

depressurization rate Refer to Vega User’s manual no




Thermal Environment Fluxes before and after Fairing jettisoning

Aerothermal and Radiative fluxes

Refer to Vega User’s manual yes

EMC Environment Spurious radiation interference levels from LV

Refer to Vega User’s manual no

Cleanliness Cleanliness Refer to Vega User’s manual no 5.5. S/C qualification Qualification to random: The spacecraft structure shall demonstrate compliance with the random vibration environment in the 20Hz - 2000Hz frequency range (refer to the levels in the above table section 5.4). Qualification to shock: When using the SSMS dispenser, the sizing shock event is the S/C separation event. The spacecraft shall demonstrate to be able to stand this separation shock.




6. Stack integration & Launch campaign This chapter describes the logic for the integration of the Stack and the activities during Launch campaign. The stack integration will be done in a dedicated integration facility in Europe, with the exception of the Mini Satellite (Class#1) for which the usual processing flow will apply (refer to Vega User’s Manual Section 6). The overall logistic scheme is shown below:

Spacecraft Class#2,3,4

Manufacturer

Lightband

PSC

SSMS dispenser

SAB SSMS Stack Integration

Facility

Europe

Spacecraft Class#2,3,4

Manufacturer Spacecraft Class#2,3,4

Manufacturer

Spacecraft Class#1

Manufacturer

HPF / Encapsulation

Facility

CSG

Launch Pad

CSG

Spacecraft in its container Spacecraft in its container

Stack in its container

Vega Upper Composite




6.1. Stack integration flow Requirement Reference Negotiable? Notes

Micro, Nano Satellites & Cubesats integration

Micro, Nano Satellites & Cubesats mating on SSMS dispenser

In Europe, in SSMS Stack Integration facility

yes

Integration in CSG facilities available as an option to be ordered by customer

Mini Satellite integration

Mini Satellites & Cubesats dispensers mating on SSMS dispenser

At CSG (refer to Vega User’s manual Section 6)

yes

6.2. S/C checks / accessibility after Stack integration Requirement Reference Negotiable? Notes

S/C Accessibility in Europe

Operations to be performed on the S/C

Not Allowed yes To be evaluated on case by case basis.

S/C Electrical checks in Europe

By means of S/C skin connection

Not Allowed yes

To be evaluated on case by case basis. If any, the S/C skin connection shall be located on the S/C face facing the fairing.

By means of S/C-LV connectors

Allowed n/a

The number of available lines through the connectors is reported in Section 4.

S/C Accessibility at CSG

Operations to be performed on the S/C

Not Allowed yes

For any “Remove before flight” items or S/C skin connections.

S/C Electrical checks at CSG

By means of S/C skin connection

Not Allowed yes

In case of anomaly only. If any, the S/C skin connection shall be located on the S/C face facing the fairing.

By means of Dispenser/LV umbilical links

Allowed n/a

The number of available lines through the connectors is reported in Section 4.




6.3. S/C hazardous propellant loading after Stack integration Requirement Reference Negotiable? Notes

Mini Satellite (Class#1)

S/C hazardous propellant loading

Allowed At CSG, in HPF, before mating on SSMS dispenser

n/a

Micro Satellite (Class#2)


Allowed At CSG, in HPF, after Stack integration

yes

Nano Satellite (Class#3)


Not allowed

Cubesats deployer (Class#4)

- - n/a

ESA UNCLASSIFIED - For Official Use

Issue Date 13 February 2017 Ref VG-AO-0-C-30001-ESA

ANNEX 2

QUESTIONNAIRE

VEGA PROGRAMRange Uncertainties

Light Satellite NameCountry of Operator/OwnerOperator/OwnerUsers and Mission Scope e.g.: EO, SCI

Mission description for info

Planning and Current Project phase schedule to Acceptance review (Gantt Chart if available)

Planned Launch dateLaunch window year launch period, day time

Flight Lifetime years, months

Attitude control concept 3-axes, spinMission specific constraints e.g.: solar aspect angleS/C sketch in stowed state including reference frame

Class of Orbit LEO/MEO/GEOType of Orbit SSO/circular/polarfor SSO, Local Time of Ascending Node [degree]Longitude of ascending node [degree]Perigee [km]Apogee [km]Inclination [degree]Argument of Perigee [degree]Eccentricity

Launch Mass [Kg]Dry Mass [Kg]

Dimensions excluding protrusions

Protrusions positions and dimensions

C.o.G. position w.r.t. the S/C separation plane (nominal and dispersed values)wrt a ref. system centered at intersection between SC long axis and separation plane

MoI w.r.t. S/C reference coordinate system where the S/C CoG is the origin (Nominal and dispersed values)S/C Interface ring sketch, dimensions, characteristics

Longitudinal modes frequencies (all modes up to 125 Hz) and relevant participating mass (in % w.r.t. the launch mass)

FEM model if available (NASTRAN files)

Lateral modes frequencies (all modes up to 125 Hz) and relevant participating mass (in % w.r.t. the launch mass)

Preferred Separation system concept pyro/non-pyroPreferred Separation Directions wrt S/C ref. frame Lateral/LongitudinalPreferred Separation orientation constraints wrt to inertial ref. frameSeparation constraints e.g.: solar aspect angle

Needed visibility, duration for commissioning [s]Max. angular rate and delta velocity range

Total Power [W]

by S/C equipment during SB on ground and launch vehicle flight phases

Critical elements acceptance temperature limits [C] Min and Max

# of electrical links S/C - EGSE

# of electrical links S/C - Launcher

RF interface requirements

Level of cleanliness and contamination

cleanliness and related contamination constraints

Accessibility requirements/constraintsPropellant [Y/N]

Propellant identification to indicate the type of propellant useddensity of liquid [Kg/m^3]volume of tank [l]fill factorliquid volume [l]liquid mass [Kg]CoG of propellant loaded tank w.r.s. to S/C coordinate systemPressurant identification to indicate the type of pressurant used

MEOP

In flight Electromagnetic Environment to indicate the RF emission in flight

On ground Electromagnetic Environment to indicate the RF emission on ground

Ground station network requirements list of needed GS in relation to separation and acquisition phases

Launch campaign requirements specific needs during campaign, e.g.: logistics, comms, clean room provisions

CAD model if available (STEP file)

to provide the number of pins for each needed connector (Lines configuration

if available - Excel file)

Misc.

EMC

Mission Description

Thermal

characteristics

Electrical interfaces

AIVO

rbital parameters

PropulsionSeparation

Physical characteristicsM

ech. Prop.

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