VEGA. 1.7 Launch cost : 18.5 M$ (target price in mature production phase for ESA and European government customers)
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1 1. IDENTIFICATION 1.1 Name 1.2 Classification Family : Series : Version : Category : SPACE LAUNCH VEHICLE Class : Small Launch Vehicle (SLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : ELV-S.p.A (1) Corso Garibaldi Colleferro, Roma, ITALY Telephone: Fax: Development manager : ESA (European Space Agency) 8-10, rue Mario-Nikis Paris, FRANCE Telephone: 33(1) Fax: 33(1) Vehicle operator : ARIANESPACE Boulevard de l'europe - BP Evry Cedex, France Telephone: 33(1) Fax: 33(1) 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 : First launch : Planned for end 2007 (1) ELV-S.p.A. is an AVIO S.p.A / ASI joint venture December 2004 Page 1
2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability Low Earth Orbits The 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 LV PERFORMANCE REQUIREMENT Payload Mass [kg] km 500 km 700 km 1200 km 1500 km Orbit Inclination [ ] FIGURE 1 - LOW EARTH ORBIT PERFORMANCE Geosynchronous and Interplanetary Orbits No capability Injection accuracy The 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. December 2004 Page 2
3 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 4 th stage is sequenced to carry out a manoeuvre to avoid subsequent collisions Spin-up performance The roll control system can provide a spin rate 5 r.p.m. clockwise or counter clockwise 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 /s Spinning < 5 (nutation) +/- 1 /s 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 Payload compartments and adaptors 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 December 2004 Page 3
4 Multiple launch capability: single launch: multiple launches 1 main + 6 micro-satellites 2 main payloads in the kg range The S/C is mounted, using a clampband, on top of a conical standard adaptor ( 1920 mm on base mm on top) The configuration can be: - A main S/C mounted on top of the adaptor and up to 3 microsatellites 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 Fairing Overall length: 7.88 m (3.5 m cylindrical part) Volume: m 3 Usable volume (static + dynamic): see Figure 3 (usable diameter: 2.35 m). FIGURE 3 - FAIRING EXTERNAL DIMENSIONS AND PAYLOAD VOLUME December 2004 Page 4
5 3.4 Environments Mechanical environment Spacecraft design and dimensioning data: - Spacecraft materials: Total mass loss: 1 % Volatile condensable material: 0.1 % - Dimensionning: Spacecraft modes: Lateral frequencies: Longitudinal frequencies: Allowable masses, COG positions and inertia: 15 Hz 35 Hz Mass (kg) Centre of gravity (mm) (from separation plane) SINGLE LAUNCH DUAL LAUNCH MAIN PAYLOAD AND MICRO-SATELLITES Upper: Lower: Main: Micro-satellites (max 6): < Static unbalance (distance d between COG of payload and the L/V roll axis Dynamic unbalance (angle ε between principal axes of inertia and L/V roll axis) Spinned payloads: d 15mm 3 axes controlled payloads: d 30 mm Spinned payloads: ε 1 3 axes controlled payloads: ε 6 Will be defined by Arianespace Will be defined by Arianespace Will be defined by Arianespace Will be defined by Arianespace 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 December 2004 Page 5
6 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 ± 1.50 ± 1.50 P 80 flight ± 1.00 ± 1.00 Z 23 flight ± 1.00 ± 1.00 Z 9 flight ± 1.00 ± 1.00 Z 9 / 23 Ignition (transient) /-5.50 ± 0.50 Low frequency vibrations: TABLE 1 - FLIGHT LIMIT LOADS - 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 [ ] Hz (see Figure 4) at the LV/PL interface. 0.1 RMS LEVEL[g] = 5 PSD (g2/hz) g2/hz Frequency (Hz) FIGURE 4 - RANDOM VIBRATIONS LEVELS December 2004 Page 6
7 3.4.2 Acoustic vibrations The envelope spectrum of the noise induced inside the fairing during flight is shown in Figure Octave Center Frequency [hz] Flight Limit Level (db) Noise Level [db] Octave band - Frequency [Hz] FIGURE 5 - ENVELOPE OF ACOUSTIC SPECTRUM OASPL ( Hz) 142 OASPL: Overall Acoustic Sound Pressure Level 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=10 Radial and axial Maximum envelope (on S/C ring side) SRS [g] 100 Freq. SRS [Hz] [g] freq [Hz] FIGURE 6 - MAXIMUM SHOCK RESPONSE SPECTRUM (Q = 10) ON THE S/C INTERFACE RING WITH THE L/V IN BOTH AXIAL AND RADIAL DIRECTIONS December 2004 Page 7
8 3.4.4 Thermal environment Net flux radiated inside the fairing: Aerothermal fluxes at fairing jettisoning: 1000 W/m² 1135 W/m² Variation of static pressure Maximum slope during ascent phase: 50 mbar/s Maximum pressure differential over ambient at fairing separation: 80 mbar Contamination and cleanliness Organic deposits on the spacecraft (ground operations): Organic deposit on the spacecraft (flight): Obscuration factor induced by launch activities: Cleanliness conditions : 2 mg/m²/week 2 mg/m² < 500 ppm S/C location Transfer between buildings S/C in EPCU Transfer to launch zone S/C on L/V In CCU container Not encapsulated Encapsulated Encapsulated On launch pad Cleanliness class Radio and electromagnetic environment RADIATION CHARACTERISTICS (SEE TABLE 2) Telemetry Telecommand-Destruction reception Trajectography transponder MHz / 8 W MHz 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: GHz 145 dbµv/m GHz GHz 35 dbµv/m 2.2 GHz GHz 145 dbµv/m 2.9 GHz GHz 145 dbµv/m 5.4 GHz GHz 145 dbµv/m GHz GHz 45 dbµv/m 14 GHz GHz 55 dbµv/m December 2004 Page 8
9 SPURIOUS RADIATION INTERFERENCE FROM THE PAYLOAD (SEE FIGURE 8) 14 khz - 1 GHz 100 dbµv/m 420 MHz MHz 35 dbµv/m 2 GHz - 18 GHz 145 dbµv/m 5.45 GHz 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) Transmitters Receivers Radar transponder system (RT) Radio-destruct receiver (RTX) 3 rd stage/avum interstage AVUM Transmitters Telemetry transmitter 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 December 2004 Page 9
10 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 December 2004 Page 10
11 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 December 2004 Page 11
12 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 operations can be divided in 6 main phases: - Pre-campaign (to be completed before the campaign itself): 1 st 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 December 2004 Page 12
13 4. LAUNCH INFORMATION 4.1 Launch site 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 ( to ), 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 - General description FIGURE 10 - EUROPE SPACEPORT 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, 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 launcher. December 2004 Page 13
14 Launch Complex FIGURE 11 ARIANE/ 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 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 Control Centre. December 2004 Page 14
15 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) 0 s s s 211 s 216 s 322 s 356 s 699 s s s s s EVENTS 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 December 2004 Page 15
16 4.3 Launch record data None Provisional reliability: Planned launches First demonstration launch scheduled end 2007 Stabilized expected launch rate: four launches per year 5. PRESENTATION 5.1 Launch vehicle Payload Fairing AVUM 4 th stage 3 rd stage Z 9 SRM 2 nd stage Z 23 SRM 1 st stage P80 SRM FIGURE 13 LAUNCH VEHICLE December 2004 Page 16
17 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass : 30.2 m : 3 m : 137 t 5.3 General characteristics of the stages STAGE 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) Diameter (m) Dry mass (t) Propellant: Type Solid Solid Solid Liquid Fuel HTPB 69/19/12 HTPB 69/19/12 HTPB 69/19/12 UDMH Oxidizer N 2 O 4 Propellant Mass (t) Fuel Oxidizer Tank pressure (bar) ,6 Total lift-off mass (t) (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 December 2004 Page 17
18 5.4 Propulsion STAGE 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 Engine mass (t) Engine length (m) Engine diameter (m) Feed system type N/A N/A N/A Pressure-fed Chamber pressure (bar) Cooling Ablative Ablative Ablative - Mixture ratio N/A N/A N/A - Specific impulse (t) Sea level Vacuum Thrust (kn) Sea level Vacuum Burning time (s) Up to 667 Nozzle expansion ratio Restart capability N/A N/A N/A 5 (1) AVIO S.p.A - SNECMA Moteurs joint venture FIGURE 16 - ZEFIRO 9C FIGURE 17 - AVUM FIGURE14 - P 80 FIGURE15 - ZEFIRO 23 December 2004 Page 18
19 5.5 Guidance and control Inertial Measurement Unit (IMU) guidance system STAGE Pitch, yaw, roll TVC Electro-actuator nozzle gimballing TVC Electro-actuator nozzle gimballing TVC Electro-actuator nozzle gimballing + Roll control by AVUM upper stage TVC Electro-actuator nozzle gimballing + Roll control by AUM GN2 cold gas ACS Max deflection DATA SOURCE REFERENCES 1 - User s Manual Issue 2 September ESA - Reaching for the skies - n 23 - September ESA internal sources 5 - An overview of small launch vehicle propulsion - Europropulsion, FiatAvio Presentation AAAF Symposium - Propulsion for Space Transportation in the XXIst Century Versailles, France May, System Design Review Data package December 2004 Page 19
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