[american institute of aeronautics and astronautics 40th aiaa/asme/sae/asee joint propulsion...

American Institute of Aeronautics and Astronautics1

VEGA LAUNCH VEHICLE PROPULSION SYSTEMS – AN OVERVIEW OF THE 2004 DEVELOPMENT STATUS

Stefano BianchiESA - Vega Launch Vehicle Manager

ESRIN - Frascati (Rome) – Italy

Michel BonnetESA – Head of the Vega StagesESRIN - Frascati (Rome) – Italy

Alessandro TrippiESA – Liquid Propulsion EngineerESRIN - Frascati (Rome) – Italy

Agostino NeriAvio S.p.A – Product Engineering, Space Solid Propulsion Chief Engineer

Colleferro (Rome) – Italy

Roberto FabriziELV – Vega Programme Manager

Colleferro (Rome) - Italy

ABSTRACT

This paper presents a technical and programmatic overview of the Vega launcher, with a specific focus on its propulsion system. Vega is the new European Small Launch Vehicle developed by ESA.The propulsion system of the Vega LV is composed of three solid rocket motors (monolithic boosters), namely P80, Zefiro 23 and Zefiro 9, a bi-propellant (UDMH/NTO) liquid stage for apogee boost and orbit circularization and a cold gas (GN2) Attitude and Control System.The overview starts with the description of the main characteristics of the Vega launcher and gives the current status of its development. System requirements and optimization analysis have led to the current allocation to the propulsion (sub)subsystems and launcher configuration that has gone through the System Design Review.Vega SRM’s are high mass fraction, high performance motors for which new

technologies have been selected in order to comply with the performance and cost requirements. They are based on high strength CFRP material for motor case (Filament Wound) and skirts, low-density rubber for the internal insulation, low binder content and high aluminium percentage for propellant, consumable igniter, and low cost carbon/carbon material for the nozzle. The Liquid Propulsion System is mainly based on existing components and technologies. In particular, it re-uses an existing bi-propellant Ukrainian engine, and a Russian PED (polymeric bladder) tank concept.

NOTATION

ACS Attitude Control SystemASI Agenzia Spaziale ItalianaAVUM Attitude and Vernier Upper ModuleC/C Carbon-Carbon MaterialCDR Critical Design ReviewCNES Centre National d’Etude SpatialesCRFP Carbon Fiber Reinforced PlasticDM Development Model

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit11 - 14 July 2004, Fort Lauderdale, Florida

AIAA 2004-4212

Copyright © 2004 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit AIAA-2004-4212Fort Lauderdale,Florida 11-14 July 2004


EMA Electromechanical ActuatorEPDM Ethylene Propylene-Diene MonomerESA European Space AgencyFW Filament WoundGNC Guidance, Navigation and ControlGQR Ground Qualification ReviewHTPB Hydroxy-terminated PolybutadieneIPDU Integrated Power Drive UnitITE Integral Throat EntranceLPS Liquid Propulsion SystemLV Launch VehicleMEA Main Engine AVUMPEO Polar Earth OrbitP/L PayloadQM Qualification ModelSPDR System Preliminary Design ReviewSRB Solid Rocket BoosterSRM Solid Rocket MotorSSO Sun Synchronous OrbitTVC Thrust Vector ControlZefiro Zero and First Stage Rocket

INTRODUCTION

The Vega launch system is currently in development for the European Space Agency. Beginning of 2003, the main contracts covering the development and qualification of the P80 FW, the Vega first stage, and the Launch vehicle were signed between ESA and ELV, for the launch vehicle, and between CNES, on behalf of ESA, and Avio for the P80 FW, the first stage of Vega.System testing as well as manufacturing and testing of the first Solid Rocket Motor cases has successfully started in 2004. In June 2004, the milestone of the System Design Review was successfully achieved, and by the end of 2004 it is planned to complete for several launcher subsystems, the detailed design phase.The first launch of Vega is scheduled by end 2006.Vega will complete the range of launch services offered by Europe, with the objective of maintaining independent access to space, as well as a share of commercial launch service in order to make this access affordable.Launch service competitiveness and cost limitation are major drivers for the development. In this regard, key factors are the streamlining of the industrial organization and the maximum exploitation of technologies and

hardware developed within Ariane and other European national programs, production facilities and CSG (Kourou – French Guiana) launch infrastructures.On the basis of the market needs, the required in-orbit capability for the reference mission is specified as:

• 1500 kg to a 700 km altitude, circular polar orbit.

In addition to the reference mission, Vega will be able to launch satellites into orbit for a wide range of missions and applications, with a range of orbital inclinations from 5.2 degrees to SSO, altitude between 300 and 1500 km, and payload mass between 300 and 2500 kg.

VEGA LAUNCH VEHICLE DESCRIPTION

Vega is a single body vehicle composed of three SRM stages, a liquid propulsion upper module, and a Fairing as shown in Figure 1.

Figure 1 – Vega Launch Vehicle

In a single launch configuration, Vega provides a minimum volume allocated to the payload is a cylindrical volume of 2.35 m diameter and 3.5 m height plus a frustum volume of 2.8 m height.The Launcher at lift-off is 30.2 meters high and weights 139 tons.The three SRM stages perform the main ascent phase while the AVUM compensates the solid propulsion performance scattering, circularizes the orbit and executes the de-orbiting manoeuvres. This module also provides roll control during the 3rd stage boost phase, and the 3-axes control during ballistic phases up to the payload separation.The typical launch sequence is shown in Figure 2.



( A l t i t u d e v s . t i m e )

P 8 0 F W I g n i t i o n& L i f t - O f fT ( m i s s i o n t i m e ) = 0 sz ( a l t i t u d e ) = 0 k mV ( r e l a t i v e s p e e d ) = 0 m / sR ( d o w n r a n g e ) = 0 k m

S t a g e 1 S e p a r a t i o nT = 1 0 7 sz = 4 4 k mV = 1 8 2 5 m / sR = 6 7 k m

Z 2 3 B u r n o u tT = 1 7 9 sz = 9 5 k mV = 4 1 2 0 m / sR = 2 7 1 k m

S t a g e 2 S e p a r a t i o nT = 2 1 7 sz = 1 1 6 k mV = 4 0 7 2 m / sR = 4 2 3 k m

S t a g e 3 S e p a r a t i o nT = 3 2 9 sz = 1 4 6 k mV = 7 6 4 5 m / sR = 1 0 4 5 k m

T r a n s f e r O r b i tI n j e c t i o nT = 7 2 5 sz = 2 1 6 k mV = 7 9 1 2 m / sR = 4 0 3 1 k m

V E G A R e f e r e n c e T r a j e c t o r y

( 7 0 0 k m , P E O )

F a i r i n g S e p a r a t i o nT = 2 2 3 sz = 1 1 8 k mV = 4 1 6 2 m / sR = 4 4 5 k m

Figure 2 – Vega Launch Sequence

The performance map is reported in Figure 3.

1000

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2300

0 10 20 30 40 50 60 70 80 90

Orbit Inclination [°]

Pay

load

Mas

s [k

g]

VEGA LV PERFORMANCEREQUIREMENT

Figure 3 – Performance Map

SOLID ROCKET MOTOR TECHNOLOGIES

The three VEGA SRM stages use several common concepts and technology:• CFRP case, with integrated skirts, obtained

by filament wounding,• Low density (internal) thermal protection,• Igniters with consumable case,• Low torque flexible joint, • TVC with electromechanical actuators

(EMA).Moreover the Vega SRM benefit of the experience gained in the development and

qualification of the Zefiro 16 tons solid rocket motor.

Three firing tests of the Zefiro 16 motors have been successfully performed in the years 1997-2000.

Figure 4 – Filament winding process

Motor case and skirts

SRM cases are manufactured by winding helical and hoop layers of a prepreg tow onto a metallic mandrel already covered with the thermal insulation (Figure 4).



The skirts are laid up by prepreg tape plies, and co-cured with the pressure vessel in order to guarantee a high strength skirt-to-vessel connection. The winding mandrel is then dismounted.End rings, bonded and bolted together with the skirts, provide interface to interstages structures.The choice of the Carbon fibre and resin system is the result of a trade-off driven by mechanical strength, glass transition temperature of the resin, perenniality, recurring cost and material handling. Figure 5 shows the filament winding of the TM model of the Zefiro 23 motor.

Figure 5 – Winding of the Zefiro 23 TM

M30S fibre was chosen for the P80 case but T1000G for Zefiro 23 and Zefiro 9 because of the stringent requirements on the inert masses. The same resin system - UF3325 -is used for all three cases.

Thermal insulation

The internal thermal protection is a low-density EPDM rubber filled with aramid fibres and glass micro-spheres. This formulation has been tested on Zefiro 16 DM2 and QM1 models, demonstrating a very good ablative and mechanical resistance to the aerothermodynamic fluxes exerted by such type of motor.The technology for the application of the thermal insulation is an improved semi-automatic tape wrapping called PTS: the rubber tape, supported with an aramid roving, is wrapped onto the filament winding mandrel

after positioning the insulated polar bosses and stress relief flaps (Fig. 5 and 6).

Figure 6 – PTS Application Process

The size of the tape can be tailored thus optimize the smoothness of the insulation surface, in particular in domes area.

Figure 7 – PTS application

Propellant grain and liner

The propellant is a HTPB 1912 formulation with 19% of aluminium and 12% of binder. Tri-modal ammonium perchlorate distributions have been preferred for Zefiro 23 and Zefiro 9 for the expected reduction of alumina slag deposit as demonstrated by Zefiro 16 firing tests results, while bi-modal distribution has been retained for the P80 essentially for cost reduction. Both US (Wecco) and European (SME) sources are available and compatible with the required performances.

The P80 SRM will use a RV4 HTPB polymer, whose production in Ravenna plant, in Italy, has been recently qualified. RV4 may also substitute in the short-term the R45HT (Atochem) used on Ariane 5. A strain-



improved grade, the RV4-2, is needed for Zefiro 23 and Zefiro 9, whose grain deformation is higher than P80 one.The AP grinding and Iron Oxide content is tuned so that to comply with the required burning rates: 10.2 mm/s for P80, 8.2 mm/s for Z23 and 6.2 mm/s for Z9.The geometry of the propellant grain has been selected in order to provide a burning surface evolution according to LV thrust requirements. A finocyl configuration, consisting of a forward cylindrical region and an aft star region (11 tips for P80 and Z23; 9 tips for Z9) has been chosen providing at the same time grain tailor-ability and high volume fraction. In all cases the star region is positioned in the aft of the motor, thus simplifying the decomposable casting mandrel design and the relevant fin extraction device.

Igniter assembly

A two-stage igniter architecture has been selected. The pyrotechnic igniter is charged with BKNO3 pellets. The main igniter is basically composed by (Fig. 8):• Adapter ring, realized of high strength

aluminium alloy forging, insulated with EPDM rubber;

• A consumable carbon epoxy composite case;

• A star-shaped propellant grain, realized with HTPB1414.

Several nozzles, machined in the dome case, allow for the hot gas flow impingement to the motor propellant grain.

Figure 8 – Igniter AssemblyNozzle assembly

The basic design concept of nozzles is a classical submerged downstream pivot point. The design aims to simplify and reduce the number of parts and interfaces in order to reduce the manufacturing time and to improve the robustness of the assembly.

Figure 9 - P80 Nozzle Assembly

Zefiro 23 and Zefiro 9 nozzle architecture is derived from the one qualified on Zefiro 16:• Flexible joint composed of alternate layers

of stainless steel shims and low-modulus rubber pads protected by boot-straps and baffles;

• Throat made from 4D C/C Throat Insert and 3D aft Throat Divergent;

• Nose-Cap and exit cone insulators made from Carbon-Phenolic material. The 2D C/C divergent initially foreseen for Zefiro 9 has been abandoned in order to mitigate the risk associated to a short and success-oriented development program.

• Aluminium stationary shell allowing significant mass saving with respects to steel, protected with a EPDM rubber;

• Aluminium exit cone housing.

Several new material and technologies will be validated on the P80 nozzle, with the objective to reduce drastically the production cost, in particular: • Self-protected flexible joint with glass-

epoxy shims instead of steel ones;• C/C throat made from low cost Carbon

reinforcement densified with a carbon Figure 8 – P80 Nozzle Assembly matrix (Naxeco-Pyc);

• Cowl and downstream exit cone insulators made from 3D Carbon reinforced low cost Carbon-Phenolic (Naxeco-ϕ).

A cut view of the P80 nozzle assembly is presented in Figure 9.

TVC

Electromechanical TVC has been selected for all Vega stages. Such a solution permits to simplify the architecture of the TVC increasing



the reliability and contributes to the cost reduction of the motor.

The TVC system is composed of:• Two Electro-Mechanical Actuators

(EMA);• A single electronic box including the

Power Unit and the Digital Control Module. This later module will be developed as much as possible as a common part for all Stages in order to exploit all possible synergies and assure a simplified management during theproduction phase.

• A Lithium-ion Battery Set. The chosen solution allows in particular important synergies at launch vehicle level since the same basic module can be utilized for TVC batteries of all the Stages.

• The Cable Harness needed to rely all TVC components.

The EMA of the P80 SRM will the biggest and most powerful electro-mechanical actuator designed and realized in Europe for space propulsion up to now.

THE STAGES’ PROPULSION

First stage

The first stage is based on the new P80 FW SRM developed through a parallel ESA Program.

The diameter of the first stage (P80 FW) has been set to 3 meter in order to exploit the Ariane 5 SRM casting and integration facility and tooling.

Limits on the thrust shape have been imposed in order to comply with the:• Minimum acceleration at lift-off (Region

A),• Maximum dynamic pressure (Region B),• Controllability during flight (Region B) • Maximum acceleration (Region C),• Controllability at stage separation (tail-off

thrust gradient),• Burn time (Region D).The relevant forbidden regions are depicted in Figure 10.

A

B C

DAbs (dF /dt)

F [KN]

Abs (dF /dt)

Time[s]

Figure 10 – P80 FW Thrust Limits

A synthesis of the P80 FW characteristics and performances is reported in Table 1.

Characteristics & Performances ValuesOverall length [mm] 10560Outer diameter [mm] 3000Propellant mass [Kg] 88385Inert mass [Kg] 7405Burn time [s] 107Vacuum specific impulse [s] 280Nozzle expansion ratio 16MEOP [Bar] 95

Table 1 – P80 FW SRM Characteristics & Performances

The size of the P80 FW positions it as one of the largest SRM with a filament winding CFRP motor case.Figure 11 reports the nominal vacuum thrust and pressure time histories.

P80 FW SRM - VACUUM THRUST AND PRESSURE PROFILES

0

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3000

3500

0 20 40 60 80 100 120

Time (s)

Th

rust

0

20

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60

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100

Pre

ssur

e

Thrust

Pressure

Figure 11 – P80 FW Vacuum Thrust and Pressure Time Histories

In addition to the P80 FW SRM, the first stage includes airframes (interstage 1/2 and interstage 0/1), raceways tunnels, ignition pyrotechnic chain, and destruction pyrotechnic chain.



The 0/1 and 1/2 interstages are “monocoque” shell structures of high strength aluminium alloy.The 1st Stage separation is performed by a pyro-zip and 6 retro-rockets mounted in the interstage 1/2 forward part. First stage assembly is shown in Figure 12.

Figure 12 – First Stage Assembly

Second stage

The second stage, designated as Zefiro 23, is based on a stretched version of Zefiro 16 SRM and loads 24 tons of propellant. The nozzle expansion ratio is 25, this value being selected in order to comply at the same time with dimensional and performance constraints.Limitations in the outer diameter of the motors were imposed in order to reuse the existing manufacturing facilities and tooling of Zefiro 16 whenever possible, and consequently minimize the investments. Thrust limits for the second stage are driven by the maximum acceleration induced on the LV. Regressive thrust shapes with appropriatethrust level and raise rates have been specified (Figure 13).

2nd STAGE SRM

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Time [s]

Vac

uu

m T

hru

st [

KN

]

Figure 13 – Zefiro 23 Thrust Gate

Main characteristics and performance of the Zefiro 23 SRM are summarized in Table 2, while the details of the vacuum thrust time histories are shown in Figure 14.


Table 2 – Zefiro 23 Characteristics and Performance

Zefiro 23 SRM - Pressure and Thrust vs Time

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Vac

uu

m T

hru

st (

kN)

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ssu

re (

bar

)

Thrust Pressure

Figure 14 – Zefiro 23 Vacuum Thrust and Pressure Time Histories

The second stage includes also raceways tunnels, TVC, ignition and destruction pyrotechnic chains. Stage assembly is shown in Figure 15.

Figure 15 – Second Stage Assembly

Third stage

The third stage propulsion is a high mass fraction SRM called Zefiro 9, with 10 tons of



propellant; as per Zefiro 23 it is strictly derived from Zefiro 16, but downsized.Thrust limits for the third stage are driven by the maximum acceleration induced on the LV. Regressive thrust shapes with appropriate thrust level and raise rates have been specified (Figure 16).Because it is an upper stage, a high nozzle expansion ratio of 56 has been selected.

3rd STAGE SRM

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Time [s]

Vac

uu

m T

hru

st [

KN

]

Figure 16 – Zefiro 9 Thrust Gate

Main characteristics and performance of the Zefiro 9 SRM are summarized in Table 3, while the details of the vacuum thrust time histories are shown in Figure 17.


Table 3 – Zefiro 9 Characteristics and Performance

Zefiro 9 SRM - Pressure & Thrust Vs Time

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Vac

uu

m T

hru

st (

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ssu

re (

bar

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Thrust Pressure

Figure 17 – Zefiro 9 Vacuum Thrust and Pressure Time Histories

The third stage includes also the interstage 2/3 - an Aluminium shell with stiffeners -, raceway tunnels, TVC, and destruction and ignition pyrotechnic chains. Firing a pyro-zip and action of 8 separation springs allows the separation 2/3.Stage assembly is shown in Figure 18.

Figure 18 – Third Stage Assembly

LIQUID PROPELLANT STAGEVEGA fourth stage, named AVUM, is made up of:• An external skirt;• The AVUM Propulsion Module (APM)

where most of the propulsion equipment is located;

• The AVUM Avionics Module (AAM) supporting most of the LV avionics equipment,

as shown in Figure 19.



Figure 19 – AVUM Upper Stage

The AVUM propulsion includes a bi-propellant Liquid Propulsion System (LPS) that provides the necessary delta-velocities for reaching the final launcher orbit control, and a cold gas Attitude Control System (ACS) covering the following main functions:• Attitude recovery during second stage

separation;• Roll control during third stage flight;• Attitude control during coasting phases;• Orbit control for the Collision Avoidance

Manoeuvre;• Payload pointing manoeuvre.

The main characteristics of the AVUM upper stage are presented in the Table 4.

Characteristics & Performances ValuesAVUM stage dry mass [Kg] 440Propellant loading [Kg] 550Pressurant (GHe) gas loading [Kg] 3.7Main engine thrust [N] 2450LPS total impulse [kN.s] 1634Restart capability 5

Table 4 – AVUM Stage Main Characteristics

The LPS layout is shown in the Figure 20.

The LPS is a bi-propellant pressure fed system using NTO and UDMH as propellant.

The pressurant gas (Helium) is stored in one tank at the initial pressure of 310 bar and the pressurization pressure is regulated at 32.2 bar.The propellant is stored inside four bladder tanks (two tanks for the NTO and two for the UDMH). The propellant tanks are pressurized at a pad pressure of 6 bar during the ground operations. Priming of the propellant lines and pressurization of the propellant tanks is performed during the flight of the third stage.Two Non-Return Valves are located on each branch downstream the propellant tanks to prevent propellant vapour mixing during the coasting phases.

Figure 20 – LPS Layout

Two normally closed pyro-valves are located downstream the propellant tanks to isolate them during the ground operations. The two valves configuration has been selected to accommodate the large mass flow rate encountered during the main engine operation and to be compatible with existing off-the-shelf equipment.



The AVUM Main Engine Assembly (MEA) is based on the combustion chamber of the Yuzhnoye RD-869 engine. The main performance of the MEA is presented in the Table 5 while the engine is shown in the Figure 21.

Characteristics & Performances ValuesThrust [kN] 2450Specific Impulse [s] 315.5Mixture ratio (O/F) 2Inlet Pressure [bar] 30Combustion chamber pressure [bar] 20Single firing duration [s] 608Total firing time [s] 667

Table 5 – Main Engine performance at nominal operating conditions

Figure 21 – AVUM Main Engine

The engine benefits from the large heritage gathered by Yuzhnoye through an extensive development test programme (136 firing tests on 58 engine) and 158 flights on ICBM’s. The layout of the engine combustion chamber is presented schematically in the Figure 22.

Figure 22 – Main Engine layout

The combustion chamber is cooled regeneratively by the two propellants and by two film cooling rings located downstream the injector plate and upstream the throat. Temperature of the lower part of the nozzle is limited by radiative heat exchange. Two solenoid valves are used to control the propellant flow. A Cut-off-Valve is placed between the Flow Control Valve and the engine injector on the Oxidizer feeding line. This valve closes when the pressure drops below a pre-set value to reduce the engine shut down phase. The MEA operating boxes for the VEGA mission are presented in the Figure 23.



27

28

29

30

31

32

33

34

27 28 29 30 31 32 33 34

PinOx (bar)

Pin

Fu

(b

ar)

MEA operating keypoint at 98%expulsion leveldegradeted mode(99.5% EOL)

operating box at 3sigma

"MEA Qual box"

"MEA develop box

Figure 23 – MEA operating boxes

The propellant bladder (PED) tank is shown schematically in the Figure 24 while the tank main characteristics are presented in the Table 6. The Teflon bladder is derived from the technology already developed by Babakin Space Centre and flown on Phobos-88 and Mars-96 spacecraft and on the Fregat upper stage. A Titanium shell is being developed to reduce the tank mass.

Figure 24 – Bladder (PED) tank

Characteristics & Performances ValuesMinimum volume [litre] 142Propellants Oxidiser Fuel

NTOUDMH

Max Ox loading [Kg] 367Max Fu loading [Kg] 183Expulsion efficiency [%] 99.5MEOP [bar] 35.6Proof/Burst factors 1.25/1.5Dry mass [Kg] 16.2

Table 6 – Propellant tanks characteristics

The characteristics of the pressurant gas tank are given in the Table 7. The same type of tank is also used to store the Nitrogen gas used by the ACS.

Characteristics & Performances ValuesVolume [litre] 87MEOP [bar] 310Proof/Burst factors 1.5/2.0

Table 7 – Gas Tanks main characteristics

The ACS layout is presented in the Figure 25.



Figure 25 – ACS Layout

The Nitrogen gas is stored in one single gas tank and feed two clusters of thrusters. A single stage pressure regulator is used to reduce the pressure from the gas tank pressure (310 bar at the beginning of mission) to the operating pressure of the thrusters, 72.6 bar.Two thruster clusters are located on the opposite side of the AVUM skirt. The MOOG 50-820 model has been adopted. Three thrusters are grouped in each cluster. The performance of the ACS thrusters is presented in the Table 8. As stated above, the same model of tank is used for the ACS gas and the LPS pressurant gas.

Characteristics & Performances ValuesNominal thrust level [N] 50Inlet pressure [bar] 72.6ACS Gas loading [Kg] 17MIB [N.s] 0.5

Table 8 – ACS Thruster Characteristics

DEVELOPMENT PLANS

SOLID PROPELLANT MOTORSThe Development and Qualification Plans of all the three SRMs have commonalties as far as the logic and the number of specimens/tests. The key elements of the development logic are essentially:• To qualify the motor case in terms of

design, technologies and manufacturing. This is obtained with the TM and DM0 motor cases. In particular, these models will undergo the hydro-proof/burst tests and mechanical tests using the loads acting on the LV during the operational mission.

• To perform an inert casting before the first active development motor, utilizing the TM motor case.

• To perform 3 firing tests of the pyro-igniter (PY1 to PY3) and 3 tests at igniter level (DI1 to 3).

• To qualify the flexible joint completely by representative heavy actuation tests;

• To validate and qualify the TVC subsystem(s) on a functional mock-up (HWIL).

• To perform specific small and intermediatescale tests for P80 nozzle new materials behaviour characterization;

• To achieve the SRM ground qualification through two static firing tests (DM1 and QM1).

LIQUID PROPULSION SYSTEMThe LPS and ACS development and qualification programme is based on two sub-system models: the Propulsion Validation Model (PVM) and the UCFIRE model.The main objectives of the PVM tests are:• To validate the sub-system flow

characteristics (mass flow rates and pressure losses, pressure regulator characteristics etc.) against the results of the associated analytical models;

• To verify pressure peaks during the sub-system transient operations such as the LPS priming sequence;

• To validate the sub-systems operation sequence including, for the LPS, the propellant tank pressurization and the priming of the propellant lines.

The PVM tests will be performed using propellant simulants for the LPS. The sub-



system will be mostly in the flight configuration except for the main engine, which is replaced by calibrated orifices to simulate the engine flow characteristics. The PVM tests are the qualification tests for the ACS and therefore the subsystem will be in full flight configuration and the tests will be performed using Nitrogen. As for the UCFIRE, the tests will be performed on the LPS in flight configuration using both NTO and UDMH to feed the main engine. The main objectives of the tests are as follows:• To validate the sub-system flow

characteristics using the flight propellants.• To verify the sub-system characteristics

against instabilities due to coupling between the main engine and the propellant feeding lines.

• To validate the POGO models of the AVUM stage.

Several tests will also performed at equipment level. In particular three models of the Main Engine Assembly will be tested by Yuzhnoye to validate the engine performance against the VEGA requirements. To this end, two models will be subject to a full qualification programme.As for the propellant tank, several tests will be performed on the bladder material, on the whole bladder and on two tanks which will undergo a complete qualification programme.

MASTER PLANNING & PROGRESS OF

ACTIVITIES

The master planning of the Vega Propulsion Systems development and qualification program is illustrated in Figure 26.

Figure 26 – Propulsion Systems Master Planning

Preliminary Design Reviews for the three SRMs and LPS/ACS and their main components have been completed.The P80 inert casting was carried out in May 2004 in UPG plant in Guiana, using a metallic case, while the thermal insulation manufacturing and winding of the DM0 case

was started in the fully new P80 plant, in Colleferro.The Zefiro 9 TM case went trough the hydroproof and mechanical testing phase and will go to the Inert Casting in July 2004, in Avio plants, in Colleferro – Italy. Figure 27 shows the Zefiro 9 TM.



Figure 27 - Zefiro 9 TM.

The manufacturing of Zefiro 23 TM has been completed in mid June 2004 and will follow the same sequence of tests and then Inert Loading operation within September 2004.At System level, testing activities started with the wind tunnel testing, that have been finalised within May 2004, while the small scale acoustic test, to simulate the lift off, are currently on going.The System Design Review – an intermediate key point between the PDR (held in July 2001) and CDR (planned in November 2005) - has been concluded in the beginning of June 2004.The development is entering now in the phase of the preparation for the first firing tests of the Solid Rocket Motors. The first firing test is planned for mid 2005.

ACKNOWLEDGMENTS

The authors would like to acknowledge ELV, Avio, Snecma Propulsion Solide, Stork Product Engineering, Yuzhnoye, Babakin Space Center and SABCA for providing the information to realize the present paper.

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