VEGA - TU Delft Blackboard · PDF fileEUROPE VEGA 3.2 Spacecraft orientation and separation After injection into the desired orbit, the L/V cold gas Attitude Control System (ACS) provides

EUROPE VEGA

1. IDENTIFICATION

1.1 Name

VEGA

1.2 Classification

Family : VEGA Series : VEGA Version : VEGA

Category : SPACE LAUNCH VEHICLE Class : Small Launch Vehicle (SLV) Type : Expendable Launch Vehicle (ELV)

1.3 Manufacturer : ELV-S.p.A (1) Corso Garibaldi 22 00034 Colleferro, Roma, ITALY Telephone: 39 06 97285 111 Fax: 39 06 97285 648

1.4 Development manager : ESA (European Space Agency) 8-10, rue Mario-Nikis 75738 Paris, FRANCE Telephone: 33(1)5369 7388 Fax: 33(1)5369 7678

1.5 Vehicle operator : ARIANESPACE Boulevard de l'Europe - BP 177 91006 Evry Cedex, France Telephone: 33(1)6087 6000 Fax: 33(1)6087 6304

1.6 Launch service agency : ESA (European Space Agency)

1.7 Launch cost : 18.5 M$ (target price in mature production phase for ESA and European government customers)

1.8 Development cost : 430 M€ (including 40 M€ for new P80 motor and 40 M€ for ground segment)

2. STATUS

2.1 Vehicle status : Under development

2.2 Development period : 2001 – 2007

2.3 First launch : Planned for end 2007

(1) ELV-S.p.A. is an AVIO S.p.A / ASI joint venture

December 2004 Page 1

EUROPE VEGA

3. PAYLOAD CAPABILITY AND CONSTRAINTS

3.1 Payload capability

3.1.1 Low Earth Orbits

The VEGA Launch Vehicle is capable of several missions in circular orbits ranging from:

Altitude: 300 km to 1500 km

Inclinations from 5.2° to Sun Synchronous

Launch Vehicle design mission: 1500 kg Payload at 700x700 km polar.

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

0 10 20 30 40 50 60 70 80 90 100Orbit Inclination [°]

Payl

oad

Mas

s [k

g]

300 km

1500 km

700 km

1200 km

500 km

110

VEGA LV PERFORMANCEREQUIREMENT

FIGURE 1 - VEGA LOW EARTH ORBIT PERFORMANCE

3.1.2 Geosynchronous and Interplanetary Orbits

No capability

3.1.3 Injection accuracy

The VEGA injection accuracy standard (1 σ value) at 700 km altitude is better than: 10 km on altitude, ± 0.05° on inclination, ± 0.1° on ascending node.


EUROPE VEGA

3.2 Spacecraft orientation and separation

After injection into the desired orbit, the L/V cold gas Attitude Control System (ACS) provides the required orientation and spin to the spacecraft (S/C) before its separation. After completion of the separation, the 4th stage is sequenced to carry out a manoeuvre to avoid subsequent collisions.

3.2.1 Spin-up performance

The roll control system can provide a spin rate ≤ 5 r.p.m. clockwise or counter clockwise.

3.2.2 Spacecraft pointing accuracy

Pointing accuracy after separation, for a standard payload of 1500 kg, for three axis or spin stabilised conditions, at 99% probability level:

ANGLE RATE

Yaw +/- 1° +/-0.6°/s

3-Axis Pitch +/- 1° +/- 0.6°/s

Roll +/- 1.5° +/- 1°/s

Spinning < 5° (nutation) +/- 1°/s

3.2.3 Separation velocities

Relative velocity between two separate bodies: ≥ 0.5 m/s

Typical velocity for 1500 kg payload (single launch configuration): 1 m/s

3.3 Payload interfaces

3.3.1 Payload compartments and adaptors

3.3.1.1 Adaptor

The Adapter 937 is bolted on the upper part of the Propulsion Module during Upper Composite integration with launch vehicle.

FIGURE 2 - 3RD STAGE LAYOUT


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Multiple launch capability: 1 main + 6 micro-satellites 2 main payloads in the 300-1 000 kg range

single launch: The S/C is mounted, using a clampband, on top of a conical standard adaptor (∅ 1920 mm on base - ∅ 937 mm on top)

multiple launches The configuration can be:

- A main S/C mounted on top of the adaptor and up to 3 micro-satellites mounted on a platform (like the ARIANE Structure for Auxiliary Payload - ASAP concept), fixed on the adaptor,

- A satellite cluster laid on a dispenser, mounted on top of the adaptor

Minimum separation velocity: 0.5 m/s

3.3.1.2 Fairing

Overall length: 7.88 m (3.5 m cylindrical part) Volume: 19.68 m3 Usable volume (static + dynamic): see Figure 3 (usable diameter: 2.35 m).

FIGURE 3 - FAIRING EXTERNAL DIMENSIONS AND PAYLOAD VOLUME


EUROPE VEGA

3.4 Environments

3.4.1 Mechanical environment

Spacecraft design and dimensioning data:

- Spacecraft materials:

∗ Total mass loss: ≤ 1 % ∗ Volatile condensable material: ≤ 0.1 %

- Dimensionning:

Spacecraft modes:

∗ Lateral frequencies: ≥ 15 Hz ∗ Longitudinal frequencies: ≥ 35 Hz Allowable masses, COG positions and inertia:

SINGLE LAUNCH DUAL LAUNCH MAIN PAYLOAD AND MICRO-SATELLITES

Mass (kg) 300-2500 Upper: 300-1000 Lower: 300-1000

Main: 300-2000 Micro-satellites

(max 6): < 100

Centre of gravity (mm) (from separation plane) 200-2000 200-1700 ≤ 450

Static unbalance (distance “d” between COG of payload and the L/V roll axis

Spinned payloads: d ≤ 15mm 3 axes controlled payloads:

d ≤ 30 mm

Will be defined by Arianespace


Dynamic unbalance (angle “ε” between principal axes of inertia and L/V roll axis)

Spinned payloads: ε ≤ 1° 3 axes controlled payloads: ε ≤ 6°



Roll inertia “Ir”

≤ 10 m².kg

Pitch “It” and Yaw inertia

≤ 10 m².kg

λ= It / Ire Spinned payloads: λ > 1 3 axes controlled payloads:

0.4< λ<2.5 N/A N/A


EUROPE VEGA

Steady state acceleration and quasi-static loads:

- Peak acceleration for a payload above 600 kg: ≤ 5.5 g - Highest lateral acceleration: 1g - Flight limit loads:

Acceleration (g) Longitudinal Lateral

Flight Event Static Dynamic Static + Dynamic

Lift-off - 2.00 ± 1.50 ± 1.50

P 80 flight - 4.40 ± 1.00 ± 1.00

Z 23 flight - 6.00 ± 1.00 ± 1.00

Z 9 flight - 5.00 ± 1.00 ± 1.00

Z 9 / 23 Ignition (transient) - +3.00/-5.50 ± 0.50

TABLE 1 - FLIGHT LIMIT LOADS

Low frequency vibrations:

- Sinusoidal longitudinal vibration at the base of the S/C (2-100 Hz): < 1 g - Sinusoidal lateral vibration at the base of the S/C (2-100 Hz): < 0.8 g

Random vibrations:

The Root Mean Square (RMS) random vibration level shall not exceed 5 g in the range [20-2000] Hz (see Figure 4) at the LV/PL interface.

0.001

0.01

0.1

10 100 1000 10000

Frequency (Hz)

PSD

(g2/

Hz) 0.013 g2/Hz

RMS LEVEL[g] = 5

FIGURE 4 - RANDOM VIBRATIONS LEVELS


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3.4.2 Acoustic vibrations

The envelope spectrum of the noise induced inside the fairing during flight is shown in Figure 5.

120

125

130

135

140

10 100 1000 10000

Octave band - Frequency [Hz]

Noi

se L

evel

[dB

]

Octave Center Frequency [hz]

Flight Limit Level (dB)

31.5 128 65 130 125 134 250 135 500 134

1000 128 2000 124

OASPL (20 – 2828 Hz) 142

OASPL: Overall Acoustic Sound Pressure Level

FIGURE 5 - ENVELOPE OF ACOUSTIC SPECTRUM

3.4.3 Shocks

The envelope acceleration shock response spectrum (SRS) at the spacecraft base (computed with a Q-factor of 10) is showed on Figure 6. These levels are applied simultaneously in axial and radial directions.

SHOCK RESPONSE SPECTRUM - Q=10Radial and axial Maximum envelope (on S/C ring side)

10

100

1000

10000

100 1000 10000

freq [Hz]

SRS

[g]

Freq. SRS[Hz] [g]

100 202000 2000

10000 2000

FIGURE 6 - MAXIMUM SHOCK RESPONSE SPECTRUM (Q = 10) ON THE S/C INTERFACE RING WITH

THE L/V IN BOTH AXIAL AND RADIAL DIRECTIONS


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3.4.4 Thermal environment

Net flux radiated inside the fairing: ≤ 1000 W/m² Aerothermal fluxes at fairing jettisoning: ≤ 1135 W/m²

3.4.5 Variation of static pressure

Maximum slope during ascent phase: 50 mbar/s Maximum pressure differential over ambient at fairing separation: 80 mbar

3.4.6 Contamination and cleanliness

Organic deposits on the spacecraft (ground operations): ≤ 2 mg/m²/week Organic deposit on the spacecraft (flight): ≤ 2 mg/m² Obscuration factor induced by launch activities: < 500 ppm

Cleanliness conditions :

3.4.7 Radio and electromagnetic environment

RADIATION CHARACTERISTICS (SEE TABLE 2)

Telemetry 2200 - 2290 MHz / 8 W

Telecommand-Destruction reception 440 - 460 MHz

Trajectography transponder 5400 - 5900 MHz / 400 W peak

SPURIOUS RADIATION INTERFERENCE FROM THE L/V (SEE FIGURE 7)

14 kHz - 18 GHz 90 dBµV/m

Except in the following bandwidths:

1.00 - 1.50 GHz 145 dBµV/m

2.025 GHz - 2.11 GHz 35 dBµV/m

2.2 GHz - 2.29 GHz 145 dBµV/m

2.9 GHz - 3.4 GHz 145 dBµV/m

5.4 GHz - 5.9 GHz 145 dBµV/m

5.925 GHz - 7.075 GHz 45 dBµV/m

14 GHz - 14.8 GHz 55 dBµV/m

Transfer between buildings

S/C in EPCU Transfer to launch zone S/C on L/V

S/C location

In CCU container

Not encapsulated Encapsulated Encapsulated On launch pad

Cleanliness class 100 000 100 000 100 000 100 000 100 000 100 000


EUROPE VEGA

SPURIOUS RADIATION INTERFERENCE FROM THE PAYLOAD (SEE FIGURE 8)

14 kHz - 1 GHz 100 dBµV/m

420 MHz - 480 MHz 35 dBµV/m

2 GHz - 18 GHz 145 dBµV/m

5.45 GHz – 5.825 GHz 70 dBµV/m

The basic RF characteristics of the L/V transmission and reception equipment are given in Table 2.

Equipment Frequency (MHz)

Power (W) Power (dBm) Antenna (Number)

3rd stage/AVUM interstage

Transmitters Radar transponder system (RT) 5400-5900 - -70 1

Receivers Radio-destruct receiver (RTX) 440-460 - -110 1

AVUM

Transmitters Telemetry transmitter 2200-2290 8 - 1

TABLE 2 - L/V RF SYSTEM CHARACTERISTICS

The intensity of the electrical field generated by spurious or intentional emissions from the launch vehicle and the range RF systems do not exceed those given in Figure 7. These levels are measured at adapter/AVUM interface.

FIGURE 7 - INTENSITY OF ELECTRICAL FIELD GENERATED BY SPURIOUS OR INTENTIONAL

EMISSION FROM THE LAUNCH VEHICLE AND THE RANGE


EUROPE VEGA

To prevent the impact of spacecraft RF emission on the proper functioning of the L/V electronic components and RF systems during ground operations and in flight, the spacecraft should be designed to respect the L/V susceptibility levels given in Figure 8.

FIGURE 8 - L/V SUSCEPTIBILITY LEVEL AT SPACECRAFT/ADAPTER INTERFACE

3.5 Operation constraints

Launch constraints:

- Launch windows: The Vega L/V can be launched any day of the year, any time of the day respecting the specified lift-off time.

- Launch slot: 1 month within the launch period (stabilized conditions)

- Launch postponement: by 24 hours or 48 hours


EUROPE VEGA

A typical schedule for non-recurring missions is based on a 24-months timeline as shown in Figure 9. This planning can be reduced for recurrent spacecraft, taken into account the heritage of previous similar flights, or in case of the existence of a compatibility agreement between the spacecraft platform and the launch system.

Note: and - the deliverables and tasks of the Customer

FIGURE 9 – TYPICAL MISSION INTEGRATION SCHEDULE


EUROPE VEGA

Flight constraints:

During the boost phase and up to separation of the payloads, no command signal can be sent to the payload or generated by the S/C onboard system.

After the boost phase and before the S/C separation, commands can be provided to the S/C.

Integration process:

The VEGA operations can be divided in 6 main phases:

- Pre-campaign (to be completed before the campaign itself): ∗ 1st stage integration in the BIP – Bâtiment d’Intégration des Propulseurs

- Launcher propulsion operation within the Mobile Gantry: ∗ P80 (Stage 1) electrical and mechanical acceptance tests, ∗ Zefiro 23 (Stage 2) integration on P80, electrical and mechanical tests, ∗ Zefiro 9 (Stage 3) integration on Z23, electrical and mechanical tests, ∗ AVUM is integrated on Z9, ∗ Launch Vehicle acceptance tests.

The corresponding functional links are plugged to the ground facilities.

- Spacecraft preparation in EPCU: ∗ S/C preparation and checkout inside non-hazardous area ∗ Transfer to hazardous area ∗ S/C hazardous operations

- Upper Assembly constitution (S/C + adaptor + fairing) in EPCU: ∗ Transfer from Airport (or Harbour) towards EPCU Buildings ∗ Adaptor unpacking and preparation ∗ Halves fairing unpacking and preparation ∗ Integration of S/C on the adaptor ∗ Integration of the 2 half fairings around the S/C ∗ Check of the Upper Assembly ∗ Preparation to transfer.

- Combined operations with L/V: ∗ Transfer of the Upper Assembly on the AVUM ∗ Integration on the AVUM ∗ Check of Upper Assembly link to the L/V and to the ground ∗ Filling of the AVUM ∗ Arming and inspections

- Final sequence: ∗ Removal of the Mobile Gantry ∗ Launch countdown ∗ Lift-off


EUROPE VEGA

4. LAUNCH INFORMATION

4.1 Launch site

VEGA launch operations are carried out by Arianespace using the former ARIANE launch site number 1 (ELA1), located at the European Spaceport in French Guiana, the Guiana Space Centre - Centre Spatial Guyanais (CSG).

The launch campaign and the flight will use the general facilities of the CSG already developed for ARIANE, the EPCU (Ensemble de Preparation des Charges Utiles, Payload Preparation Complex), the Ground Tracking Stations, will follow the Safety authorities requirements and will use the Logistics support.

The Guiana Space Centre

French Guiana's location, close to the equator, in an area outside the hurricane zone with the possibilities of launches in Northern to Eastern directions over the ocean (- 10.5° to + 93.5°), and regular air and sea connections were factors which led to the choice of Kourou for the ARIANE launch complex.

Launching aera ELA1 – ELA2 – ELA3

Technical Center CSG

FIGURE 10 - EUROPE SPACEPORT

- General description

The CSG, operational since 1968, is managed on behalf of the ESA by the Centre National d'Etudes Spatiales (CNES). An agreement between the French government and ESA defines the rights and obligations of each party with regard to ESA's launch sites and associated facilities.

Within the CSG perimeter, the following ESA facilities are namely available:

∗ ELA1, which was designed for launching ARIANE 1, 2, 3 and which was rebuilt for VEGA,

∗ the Payload Preparation Complex - Ensemble de Préparation des Charges Utiles (EPCU): with facilities made available to Users for the preparation of their spacecraft, from its arrival in Guiana up to the actual mounting of the payload on the VEGA launcher.


EUROPE VEGA

VEGA Launch Complex

FIGURE 11 – ARIANE/VEGA LAUNCH COMPLEX

The EPCU consists of a number or geographically dispersed buildings:

∗ Buildings S1A and S1B are located in the CSG Technical Centre, and provide clean-room facilities for satellite preparation.

∗ Buildings S3A and S3B are assigned to satellite propellant filling operations and final integration, assembly of the satellites on the adapter, and satellite encapsulation into the VEGA nose-fairing.

∗ Building S3C is located close to S3A and S3B, and is used for monitoring and control of hazardous operations conducted in the latter.

Owned by ESA, the EPCU is operated bay the CSG for the benefit of ARIANESPACE customers. Agreements between CNES, ESA and ARIANESPACE define conditions for utilisation by ARIANESPACE of ELA1 and EPCU.

The preparation launch campaign also requires specific installations and support equipment that have to be created, including a gantry tower at the launch pad, fuel and other fluid supply lines, the VEGA Control Centre.


EUROPE VEGA

For the Eastward launches, the CSG radar, telemetry and telecommand stations are completed by the four down-range stations located at Natal in Brazil, on Ascencion Island, near Libreville in Gabon, and near Pretoria in South Africa, in order to continuously receive data on the launcher's trajectory and behaviour in flight.

4.2 Sequence of flight events

A typical flight sequence for reference mission is detailed below:

TIME AFTER LIFT-OFF (s) EVENTS

0 s 106-109 s 181-201 s

211 s 216 s 322 s 356 s 699 s

3 133 s 3 273 s 3 500 s 3 747 s

P80 FW Ignition and Lift-off P80 FW burnout, 1st stage separation, Z23 ignition Z23 burnout, 2nd stage separation, 1st coasting phase Z9 ignition Fairing jettisoning Z9 extinction 1st AVUM ignition (Perigee boost) 1st cut-off of AVUM, 2nd coasting phase (Perigee-Apogee) 2nd AVUM ignition (Apogee orbit circularisation) 2nd cut-off of AVUM, upper composite orientation Payload separation, 3rd AVUM ignition (AVUM de-orbiting) End of mission, burn, passivation

TABLE 3 - TYPICAL CHRONOLOGY FOR REFERENCE MISSION

FIGURE 12 - SEQUENCE OF FLIGHT EVENTS


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4.3 Launch record data

None

Provisional reliability: 0.98

4.4 Planned launches

First demonstration launch scheduled end 2007

Stabilized expected launch rate: four VEGA launches per year

5. PRESENTATION

5.1 Launch vehicle

PayloadFairing

AVUM 4th stage

3rd stageZ 9 SRM

2nd stageZ 23 SRM

1st stageP80 SRM

FIGURE 13 – VEGA LAUNCH VEHICLE


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5.2 Overall vehicle

Overall length : 30.2 m Maximum diameter : 3 m Lift-off mass : 137 t

5.3 General characteristics of the stages

STAGE 1 2 3 4

Designation P-80 ZEFIRO 23 ZEFIRO 9 AVUM (2)

Manufacturer EUROPROPULSION (1) AVIO S.p.A. AVIO S.p.A AVIO S.p.A

Length (m) 11.20 8.39 4.12 2.04

Diameter (m) 3.005 2.150 1.907 1.952

Dry mass (t) 7.42 1.85 0.83 0.418

Propellant:

Type Solid Solid Solid Liquid

Fuel HTPB 69/19/12 HTPB 69/19/12 HTPB 69/19/12 UDMH

Oxidizer - - - N2O4

Propellant

Mass (t) 88.3 23.9 10.1 0.55

Fuel 0.183

Oxidizer 0.367

Tank pressure (bar) - - - 35,6

Total lift-off mass (t) 95.72 25.75 10.93 0.968

(1) AVIO S.p.A - SNECMA Moteurs joint venture (2) Attitude and Vernier Upper Module (including the APM - AVUM Propulsion Module - hosting the

Propulsion Elements and the AAM - AVUM Avionics Module - dedicated to the Vehicle Equipment Bay)

Upper part

DESIGNATION VEHICLE EQUIPEMENT BAY FAIRING ADAPTOR

Manufacturer AVIO S.p.A CONTRAVES EADS CASA Espacio

Mass (kg) 150 (3) 490 kg 60 kg

(3) Included in the AVUM total dry mass


EUROPE VEGA

5.4 Propulsion

STAGE 1 2 3 4

Designation P-80 ZEFIRO 23 ZEFIRO 9 AVUM

Engine designation P80 FW Z23 Z9 RD869

Manufacturer EUROPROPULSION(1) AVIO S.p.A AVIO S.p.A YUZHNOYE

Number of engines 1 1 1 1

Engine mass (t) 87.44 25.9 10.5 0.016

Engine length (m) 7.7 7.6 3.6 -

Engine diameter (m) 3 1.9 1.9 -

Feed system type N/A N/A N/A Pressure-fed

Chamber pressure (bar) 97 95 67 20

Cooling Ablative Ablative Ablative -

Mixture ratio N/A N/A N/A -

Specific impulse (t)

Sea level - - - -

Vacuum 280 289 295 315.5

Thrust (kN)

Sea level 2621 - - -

Vacuum 2980 1196 280 2.45

Burning time (s) 106.8 71.7 109.6 Up to 667

Nozzle expansion ratio 16 25 56 102.5

Restart capability N/A N/A N/A 5

(1) AVIO S.p.A - SNECMA Moteurs joint venture

FIGURE15 - ZEFIRO 23

FIGURE14 - P 80

FIGURE 17 - AVUM

FIGURE 16 - ZEFIRO 9C


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5.5 Guidance and control

Inertial Measurement Unit (IMU) guidance system

STAGE 1 2 3 4

Pitch, yaw, roll TVC

Electro-actuator nozzle gimballing

TVC Electro-actuator

nozzle gimballing


nozzle gimballing+

Roll control by AVUM upper stage


nozzle gimballing +

Roll control by AUMGN2 cold gas ACS

Max deflection 8° 7° 6° -

6. DATA SOURCE REFERENCES

1 - VEGA User’s Manual – Issue 2 September 2004

2 - http://www.esa.int/SPECIALS/Launchers_Access_To_Space

3 - ESA - Reaching for the skies - n° 23 - September 2001

4 - ESA internal sources

5 - “An overview of VEGA small launch vehicle propulsion” - Europropulsion, FiatAvio Presentation AAAF Symposium - “Propulsion for Space Transportation in the XXIst Century” Versailles, France 14-17 May, 2002

6 - System Design Review Data package

http://www.esa.int/SPECIALS/Launchers_Access_To_Space

Documents

VEGA - TU Delft Blackboard · PDF fileEUROPE VEGA 3.2 Spacecraft orientation and separation After injection into the desired orbit, the L/V cold gas Attitude Control System (ACS) provides