The focus of the Italian PRORA-USV Program is on flight testing of a specific set of future generations launch vehicle related technologies. So emphasis is given to low-cost and easy-operation system configurations rather than full-scale, mission-sized vehicle. As a consequence, the approach emphasizes sub-scale, unmanned, autonomous flying laboratories used to test technology advancements at reduced cost and risk. The USV Program has been identified based on the belief that in the long run space access and re-entry will be guaranteed by today aviation-like vehicles (sometime called aerospaceplanes) and operations. Among others not less important, such vehicles will require innovation and maturation in three main areas: atmospheric re-entry, reusability, hypersonic flight. USV includes thus technology developments along these three directions, up to their validation either on ground and on board Flying Test Beds. The USV Program is structured in two parallel main activities:
• one devoted to develop technologies oriented to medium-long term needs • one devoted to design and realise three principal flight test beds (FTBs).
The USV experimental vehicles are essentially two-stage systems based on stratospheric balloon as first stage. This configuration represents the best compromise between vehicle performance, test objectives and program costs. The FTBs will be used to execute a number of flight experiments characterized by an incremental objective plan. The main flight experiments are defined to be cornerstones missions; they are the Dropped Transonic Flight Test (DTFT), the Sub-orbital Re-entry Test (SRT), the Hypersonic Flight Test (HFT), and the Orbital Re-entry Test (ORT). Two complementary missions are also included in the program: a Qualification Test tagged DTFT_0, and a second hypersonic mission (HFT-LP) with a liquid propulsion based vehicle.
INTRODUCTION
PRORA-USV is a Technology Development Program oriented towards the maturation of certain technologies deemed necessary for the timely launch of future generations reusable access-to-space transportation systems. The scenario in with the USV program is set can be summarized by the evolutionary system typology shown in the following table:
Time frame System Typology Today (US Space Shuttle)
Semi-reusable TSTO-VTHL
Tomorrow (2015-2020)
Fully- or more-reusable TSTO-VTHL
The Day-after-Tomorrow (2040-2050)
Reusable SSTO-HTHL
USV does not focus on any specific future launcher configuration. Rather it looks at technology advances that are supposed to be fundamental for foreseable next generations RLVs. The basic belief is that the real revolution in access-to-space cost reduction can be obtained only by pushing the development toward the realization of a today aviation-like system (SSTO-HTHL, sometime called aerospaceplanes). Identified technological areas requiring innovation and maturation are: sustained hypersonic flight, atmospheric re-entry, reusability. PRORA includes thus technology developments along these three directions up to their validation either on ground and on board Flying Test Beds. The latter may be used in different ways: as experimental models in themselves, or as vehicles on which passenger experiments can be allocated, or even as system demonstrators.
12th AIAA International Space Planes and Hypersonic Systems and Technologies15 - 19 December 2003, Norfolk, Virginia
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With the goal of an incremental test objective approach, USV has indicated in an experimental vehicle launched from a stratospheric balloon the best compromise between vehicle performance, test objectives and development costs.
THE PRORA-USV PROGRAM
TECHNOLOGICAL REQUIREMENTS
From a specific point of view, the USV final objective is to develop and validate a number of key technologies identified as representative of the needs for the future generations reusable space transportation vehicle. It is assumed that such a “future vehicle” will fly in not less than 40-50 years from now, taking off from a airport-like launch site, and operated as an actual civil airplane. Under this assumption future space transportation systems will have to satisfy many contemporary requirements of air transportation such as economy, reliability, prolonged service life and short turn around time. For this purpose, the methods and processes customary to space transportation must be combined with those of aeronautics. A merge of
aeronautic and space technologies seems thus probable and it is planned to be expressed in the vehicle configuration. An SSTO system seems today to be very pre-mature, but an easy first level analysis of the main characteristics of either TSTO-VTHL and SSTO-HTHL systems reveals that, among others not less important, both vehicles will specifically require innovation and maturation in three main areas: 1. Atmospheric Re-entry – the vehicle has to
withstand the typical large thermal loads encountered during re-enter to earth from LEO;
2. Reusability – the tendency towards an aviation-
like system translates in the reusability concept; 3. Hypersonic Flight – future space vehicles will
have to fly for large part of their mission to speed much greater than the speed of the sound, and will have to maneuver in such conditions safely, having also to “handle” the heavy thermal loads generated from the friction with the air.
Fig. 1 – USV Program Road Map
Payload Techn. Passenger Expsand 50 kg Piggy Back Exps
Balloon for all,but VEGA for ORTLaunching Stage
Semi-ReusableWinged-Body. 3 FTBs
basic LPbasicØ
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Fig. 2 – Top Level Product Breakdown Structure
The first two areas are today fully agreed upon, while the hypersonic flight capability appears to be somewhat underestimated. If it can be accepted when looking at Shuttle-like vehicles, it becomes unacceptable dealing with future TSTO and SSTO that will have to be able to really fly and freely maneuver under hypersonic conditions. The technology achievements along the mentioned main three areas will be validated either on ground and on board the Flying Test Beds.
MISSION REQUIREMENTS
The realization of a number of flying laboratories is thus the necessary complementary piece of USV. Not pretending to definitively solve the problem of full reusability, each FTB is planned to fly once or to be reused after important refurbishment. The resulting roadmap indicating the four Cornerstones Missions plus two supplementary ones is shown in Fig. 1. In the same figure, the part of the roadmap whose activities are already financed with about 70 M€ is evidenced. The same elements are reported in Fig. 2 where the subdivision of the USV Program into systems (USV_X) is evidenced. It is also shown that each system (USV_X) is composed by the corresponding the FTB_X plus the other elements necessary to execute the missions. In order to reduce costs, some system and operational design drivers have been defined that are common to almost all missions: − launch shall be made via stratospheric balloon,
that can be considered as 1st stage (except the final ORT mission, that will benefit from a VEGA launch)
− baseline Launch Base shall be the ASI station of Trapani-Milo in Sicily; optional launch base shall be defined for those missions that cannot eventually start from Milo
− landing and recovery are foreseen to be on the sea after a parachute driven descent.
The USV system shall perform the missions described here below. Dropped Transonic Flight Test (DTFT) - The balloon achieves an altitude of about 24 km, then the FTB_1 vehicle is dropped. Between 10 and 15-km the vehicle has to cope with transonic aerodynamics conditions. The main objective of this test is to have operative and technical confidence on particular aspects as − Separation from balloon and maneuvers during
the first few seconds of the mission − Capability to support and manage the transonic
conditions − Correlation of analytical results of flight
mechanics on stability, maneuverability, controllability.
− Capability to cope with the recovery phase (parachute deployment, capability to foresee and achieve the landing zone, ….).
Figure 3 shows these requirements with respect to the expected performances.
Fig. 3 – DTFT mission profile Qualification Test (DTFT-0) – The DTFT-0 mission has been conceived as a qualification mission in order to test some aspects of remarkable importance
35-km
10 - 15 km;M=1
⇓ ⇓ ⇓ ⇓
MISSION DTFT-0 DTFT SRT HFT HFT-LP ORT
SYSTEM QM USV_1 USV_2 USV_2b USV_3
EXP. VEHICLE FTB_0 FTB_1 FTB_2 FTB_2b FTB_3
CONFIGURATION Balloon_1 + FTB_0
Balloon_1 + FTB_1
Balloon_2 + FTB_2
Balloon_2 + FTB_2b VEGA + FTB_3
⇓ Cornerstone Cornerstone Missions
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for the USV_1 vehicle reliability. The qualification process shall involve the structure and recovery subsystem as well as the critical on site and in flight ascending operations. For these reasons the FTB_0 Model shall have the same characteristics of the FTB_1 vehicle in terms of design configuration, but with a lower level of complexity in terms of subsystems mounted onboard and functionalities implemented. The main objectives of the Qualification Model (FTB_0) mission are: • To test the FTB_1 vehicle structure and the
relevant subsystems functionalities to be utilized for the succeeding DTFT mission, like the recovery system.
• To prove the effectiveness of the releasing device, that will characterize the mechanical interface between the vehicle and the carrier.
• To verify the applicability of the DTFT operational concept.
• To test and validate a predefined set of DTFT mission operational procedures.
Sub-orbital Re-entry Test (SRT) - The balloon achieves a floating altitude of about 30 km; after the release from the balloon the solid booster motor pushes the FTB_2 vehicle in sub-orbital condition, up to an altitude of about 120 km. Then the vehicle starts the re-entry phase whose trajectory is optimized to maximize the heat load; thus, the vehicle achieves the maximum heat flux at about 25 km and keep a heat flux higher than 650 kW/m2 for about 15 sec. Under the assumption of radiation equilibrium wall conditions, it is expected that the wall temperature will exceed 2000°C on both the nose and wing leading edge stagnation points. Figure 4 shows these requirements with respect to the expected performances.
Fig. 4 – SRT mission profile
The main objectives of this test are: − To have operative and technical confidence on
re-entry aspects
− To provide research community with a Flying Test Bed able to test advanced materials under very severe conditions (more than 650 kW/m2 for 15 sec); as the nose and the leading edge of the vehicle can be removed and replaced with other noses, depending on the radius at the stagnation point, it is possible to experience several levels of maximum temperature.
Hypersonic Flight Test (HFT) - The balloon achieves a floating altitude of about 35-km; after the release from the balloon the solid booster motor pushes the vehicle at maximum speed, up to a maximum of Mach 7 in horizontal flight and with the requirement to keep Mach 6 at least 20 sec. The vehicle nose and wing leading edges achieve the maximum heat flux (higher than 650 kW/m2) and a corresponding temperature as high as 2000°C. Objective of this test is to have operative and technical confidence on horizontal hypersonic flight in terms of aerodynamics and flight mechanics behavior prediction.
35-km
M = 7
Fig. 5 – Hypersonic Flight Test (HFT)
Hypersonic Flight Test - Liquid Propulsion (HFT-LP) – The HFT mission will be repeated once the solid booster will have been replaced with a LOx-HC engine, under development within an ASI program. The objective of this mission is to increase the Mach 6 flight time, while flying the first European LOx-HC engine. Orbital Reentry Test (ORT) - The final USV mission will be launched from the ESA-GSC in Kourou by the ESA small launcher VEGA. The FTB_3 vehicle will be inserted in a 200 km circular orbit remaining attached to the AVUM upper stage of VEGA. After one or two orbits, the de-orbiting will be executed by the AVUM, and the system will be put on an almost typical reentry trajectory. Thanks to the proper combination of geometrical configuration (nose and wing leading edge radii) and reentry trajectory, the FTB_3 vehicle will experience heat fluxes of about 1300 kW/m2 and wall temperature around 2000°C. The program requirements referring to the Cornerstone Missions are summarized in Tab. 1.
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FLYING TEST BEDS POTENTIALITIES
Specific technologies embedded into the USV vehicles design will include high speed aerothermodynamics, advanced thermal protection system and autonomous flight operations. For what concerns aerodynamics, for example, it is stressed that USV offers the important capability to duplicate in flight a number of physical parameters typical of real launchers and reentry vehicles. On the other hand, it also duplicate what can be reproduced in a number of ground wind tunnel facilities, thus assuring the possibility to experimentally validate any extrapolation-to-flight correlation procedure. This is clearly shown by Figs. 6 through 8. FTB_2 represents also a very good tool to flight testing the laminar-turbulent transition phenomenon. It is matter of fact that the predictable natural transition point calculated on the vehicle fuselage while flying the SRT trajectory moves widely along the fuselage itself (Fig. 9). Removable wing leading edge and nose-cone will allow to test different high temperature materials; the actual mission profile and shape definition are, in fact, designed to expose these components to heat fluxes exceeding 650 kW/m2, (significantly higher than the today spacecraft) for a significant period of time.
Flight experiments will also demonstrate a number of advanced spacecraft technologies, such as non-traditional propulsion systems (hybrid or liquid air-breathers), composite cryogenic structures for tank, shape memory alloy based smart wing, health management systems.
VEHICLE DESCRIPTION
DESIGN APPROACH
The design approach for the flying test best has been defined on the basis of the general constraints that frame the program: (i) reduced development time and budget w.r.t. ambitious objectives of the Cornerstone Missions; (ii) very different flight conditions with a single type of vehicle, i.e. same external shape, same basic aerodynamic characteristics, same dry mass, similar inertia characteristics; (iii) long term plan (ORT mission within 2010). On the basis of that, the most important concepts that have applied at system level are: − Trade-off areas reduced at a minimum − Robust design approach not only in terms of
design margins but also in terms of mid-long term H/W availability, programmatic risks, decision margins.
The main “consequences” of this approach are:
Mission System Objectives Requirements
DTFT
FTB_1
To characterise in flight the transonic aerodynamic stability of the USV rockoon 2nd stages. To verify the forecast of critical phases (release, transonic instability management, deployment of recovery system…)
o M = 1 +/- 2 % o Altitude=10<h<15[km] o α >4 ° o L/D > 5
SRT
To demonstrate TPS materials capability to achieve without damages an aero thermodynamic-induced very high temperature level (>2000°) for a representative period.
o Max heat flux >650 [kW/m2],
o Rnose < 2 [cm], o t >15 [sec] o L/D > 2.5
HFT
FTB_2
To demonstrate the capability to fly in hypersonic steady conditions.
o M > 6 o Altitude < 40 [km] o t >20 [sec] o L/D > 2.5
ORT
FTB_3
To demonstrate the capability to perform an orbital mission and manage the relative orbital re-entry phase.
W .T . - G er ma ny : R W TH- A a c he n , D LRW .T . - B elg ium: V KIW .T . - F r a nc e : O N E R AW .T . - Ho lla nd : N LR , TU -D e lf tW .T . - R us s ia : IT A M , TS A G I
Fig. 6 – Viscous Effects: Comparison of USV wrt Re-entry Programs and European Wind
W .T. - Ge r many : R W T H- A ac h e n, D LRW .T. - B e lg ium : VKIW .T. - F r anc e : O NE R AW .T. - Ho llan d : NLR , T U - D e lf tW .T. - R u s sia : IT A M , TS A GI
Fig. 8 – Viscous and Rarefaction Effects: Comparison of USV w.r.t. Re-entry Programs and European Wind
Tunnels
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− Main system interactions performed among aerodynamics, flight mechanics and mechanical configuration (layout, mass properties, external shape)
− Orion-32 ATK Thiokol motor actually selected among off-the-shelf available systems with Thrust Vector Control
− Interchangeability of critical TPS items (nose cap and wing leading edges)
− Structure concept developed “around” the selected motor, with the possibility to load different-from-nominal quantity of propellant
− Mass budget developed thereof for FTB_2 has been increased of some 20% for design margin. The resulting total dry mass of FTB_2 is frozen at 1250 kg
− Vehicle designed to be intrinsically stable in as many envisaged flight conditions (Mach number, angle of attack) as possible.
AERODYNAMICS
The general shape of the vehicle has been chosen to translate the meaning and the top level objectives of USV Program, which wants to address topics concerning future generations RLV. As a matter of facts, that sort of RLV will spend most of energy during the atmospheric flight so their aerodynamic characteristics must be much better than the shorter term generation.
The main shape requirements are the following: − Radius at nose < 50-mm (10-mm nominal) − Maximum wing profile thickness < 8% − General fuselage shape as slender as possible;
according to the utilization of the Orion 32 motor, the fuselage rear part diameter is < 1-m
− Length < 8-m − Wing span < 3.5-m. The vehicle will be used as a flying test facility, so a good level of modularity is implemented: (i) the nose and the leading edge shall be removable in order to be able to test different materials at different heat load conditions; (ii) it shall be possible to install at least 2 classes of motors, namely a solid booster and a LOx-HC liquid motor. Data bases coming from other well known projects (opportunely scaled up or down to make them converge to USV configuration) and approximated methods were initially used, then integrated with detailed 3D Euler analysis complemented by viscous (boundary layer and/or Navier-Stokes) study in some strategic cases. Turbulence effects have been taken into account. Fig. 9 shows the pressure distribution on the leeside and windward surfaces of the vehicle, respectively, at
Mach Number equal to 0.5 and Angle of Attack equal to 0 deg.
Fig. 9 – Pressure distribution at Mach 0.5, AoA = 0.0 deg [Pa]
Fig. 10 shows the detail of the pressure distribution on the Vertical Tail under the same conditions.
Fig. 10 – Details relevant to the
V-Tail pressure distribution [Pa] Time-dependant calculations are currently going on to evaluate the effects of unsteadiness. The aerodynamic data base has benefited of wind tunnel experiments. In particular, the CIRA transonic PT1 tunnel has been systematically used to fully characterize FTB_1 that is planned to fly the Dropped Transonic Flight Test (DTFT). The diagrams shown in Fig. 11 give an example of the actual CFD aerodynamics data base that is focused in particular on sub-, trans-, low supersonic regimes. It is here evidenced that the center of pressure moves in those regimes between 67% and 75% of the length of the vehicles.
Fig. 11 – 3D Euler Flow Field
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AEROTHERMODYNAMICS
According to the overall logic of the USV program based on incremental objectives and complexity, each intermediate mission is somewhat down-rated with respect to the ORT mission, that is the real “flying” goal of USV. Considering that the ORT trajectory is characterized by an energy of about 25 MJ/kg, it appears clear that the first vehicles of the program (FTB_1 and FTB_2) will not be exposed to such an high energy situation. The direct consequence is that only poor aerothermodynamic effects will be encountered along the SRT mission. In fact, the 2 MJ/kg energy level of the Sub-orbital Reentry Test will not be enough to activate other than vibrational degrees of freedom. Nevertheless, it has been possible to design the vehicle shape and the trajectory in such a way as to generate very high wall heat flux and temperature conditions on few specific areas of the vehicle. The two points of FTB_2 in which wall temperature conditions will be extreme (> 2000 °C as objective) are the nose and the impingement zone of the bow shock on the wing leading edge, as shown in Fig. 12.
To evaluate the maximum thermal fluxes in these two areas the following approach has been used. An initially simplified engineering calculation has been performed along the whole SRT trajectory; then, on the heat peaking point along the trajectory a 3D Navier-Stokes computation on the nose and wing has been carried out. The radiative equilibrium wall temperatures and heat fluxes, along the re-entry path of the SRT trajectory are shown in Fig. 13 a) and b) respectively. The results indicate that a heat flux peak higher than 1000 kW/m2 is reached both on the nose and wing leading edge with corresponding temperatures higher than 2300 and 2200 °K. The 3D results (equilibrium wall radiative temperatures) on the nose, for the fully turbulent modeling, are shown in Fig. 14.
FLIGHT MECHANICS
The analysis and optimization of the trajectories has been accomplished taking into account the objectives of each of the Cornerstone Missions. Without entering a detailed discussion on the matter, it is here sufficient to underline and stress that: − because of the very specific nature of the USV
program, it is not important to define trajectories that may be representative of more realistic operational vehicles. The reason for all this is the need to take due account of specific
Fig.12 – Areas of FTB_2 under very high wall temperature conditions
Fig. 14 – 3D Nose Equilibrium Radiative Wall Temperatures (Fully Turbulent Computation)
SRT Trajectory -Rnose=1 cmSRT Trajectory -Rwle=1 cm
Fig. 05 –Nose and WLE re-entry equilibrium wall heat fluxesFig. 13 b) – Nose and Wing Leading Edge Heat Flux on SRT
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characteristics of either the FTBs and the mission requirements, plus the links among the succeeding missions;
− these links are real connections between two succeeding missions as, for example, in the case of DTFT and SRT. FTB_1 will fly the dropped transonic flight test in order to verify a number of aerodynamic and operational characteristics of FTB_2. Thus, at least a part of the DTFT trajectory must be representative of a part of the SRT trajectory.
With all this in mind, the actual nominal trajectories related the to four Cornerstone Missions are shown in Fig. 15. To enter more details on trajectories, FTB_2 flight is split into two parts. During the first path, dubbed captive flight, a stratospheric balloon accomplishes the first ascent phase of FTB_2. It allows the vehicle to fly through the densest layers of the atmosphere and, at the same time, to simplify the flight operations relevant to the staging. Due to the dimension of the vehicle and the relevant gondola, the selected balloon is one of the bigger allowable among COTS components. At an altitude of about 30-35 km the FTB_2 will be released from the balloon and will start its propelled
phase. After the motor burn-out it will follow an ascending path up to the ceiling altitude of about 120 km. During the last part of the trajectory the re-entry phase experimental conditions will be experienced as shown in Fig. 16.
Fig. 16 – SRT Trajectory
THERMAL CONTROL AND TPS
The thermal control system and TPS have been sized on the basis of the following starting constraints: − Internal temperature of 10° at the moment of the
boost in case of FTB_2 − External-structure temperature: CFRP allowed
limit of 150°; this is also the maximum temperature that the parachute liner can support.
The TPS at the various locations has been sized according with the following materials and thicknesses: − Windward: TABI, 20 mm − Leeward: FEI-650, 10 mm − Starbord: AFRSI, 10 mm − Nose and Leading edges: UHTC − Control surfaces: C-SiC. Dedicated local analyses have been carried out as for the nose-cone with an axis-symmetric transient thermal model (Fig. 17).
adiabatic wall
forced convection: hw, T0 &
radiation: A, ε, σ, Tamb
conduction
adiabatic wall
Fig. 17 – FTB_2 TPS nose axis-symmetric model
Fig. 15 – USV Reference Trajectories
0
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separationfrom balloon
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The transient analysis on a time extension of 80s, representative of the most critical part of SRT trajectory, yielded the results as in Figs. 18 and 19.
outer region: ZrB2
intermediate: C/SiC
inner: C/C
Maximum
temperatures
within solid
regions, [K]
ZrB2 2436.5
C/SiC 2224.2
C/C 1558.1
Fig. 18 – TPS most critical temperature field, nose particular
Fig. 19 – TPS temperature time history in critical points
CONFIGURATION LAYOUT
A detailed internal configuration analysis was performed to match CM location requirements for stability and control. Internal constraints were:
• ATK ORION 32 motor length • ATK ORION 32 CM time-evolution • Nose-down attitude at ditching • Parachute accommodation • Subsystems accommodation
This, along with the results from structural analysis on the components of the baseline architecture solution as well as the subsystems arrangement for FTB CM constraints, yielded to the internal configuration shown in Figs. 20 and 21 in which main structural elements are indicated as well as subsystem.
Boom
Nose bulk & nose
shell
Foam for buoyancy
Parachute
Hydraulicsystem
Avionic Bay
Avionic components
Ballastmass
FTB_0 ad-hoc subsystems
FTB_0/1 common subsystems
Boom
Nose bulk & nose
shell
Foam for buoyancy
Parachute
Hydraulicsystem
Avionic Bay
Avionic components
Ballastmass
FTB_0 ad-hoc subsystems
FTB_0/1 common subsystems
Fig. 20 – FTB_0/1 Lay-out
hydraulic system
RCS tanks
avionic bay
payload
parachute
TPS
Fig. 21 – FTB-2 Lay-out The total vehicle mass is constrained by the need to be launched by existing balloons. Even if it is possible to increase the launch mass lowering the floating altitude, the balloons allow for a maximum P/L launch mass due to structural reasons. In the case of FTB_2, the baseline Raven SF3-39.57 allows for a maximum P/L mass of 3628 kg, that includes the mass of the vehicle, the propellant, the gondola and the entire balloon chain. The mass budget of the FTB_2 is reported in Tab. 2, in which a comparison with the mass budget of FTB_1 is also given. As noticeable, apart from the about 2 t propellant needed to boost FTB_2 on the SRT mission, the imposed mass equivalence of FTB_1 w.r.t. FTB_2 requires the utilization of a 105 kg ballast mass on FTB_1.
ON-BOARD INSTRUMENTATION
Since the FTBs are conceived as flying laboratories, a specific attention has been devoted to the on board instrumentation architecture organization in order to cope both with system and experimental needs. In this perspective, from one side it has been foreseen an On Board Data Acquisition (OBDA) subsystem to collect data relevant to system housekeeping and
FTB_1 FTB_2[kg] [kg]
Structure and Mechanisms 630.0 331.0TCS & TPS 5.0 116.0GN&C 12.0 64.0On Board Data Handling 22.0 57.0Telemetry Tracking & Command 8.0 6.0Electrical Power 52.0 96.4On-Board Data Acquisition 20.0 6.5On-Board Software 0.0 0.0Harness 20.0 21.6Hydraulic System 55.0 53.2Recovery 186.0 155.0Propulsion (dry mass) n.a. 202.1Destruction Sub-system n.a. 21.6Passenger Experiment 40.0 40.0Ballast Mass 105.3 n.a.
Margin [%] 8.2 6.8
TOTAL DRY MASS 1250.0 1250.0
Propellant n.a. 1982.0
TOTAL MASS 1250.0 3232.0
MASS BUDGET
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flight parameters; from the other hand a Passenger Experiments (PEX) subsystem has been drawn to take into account more specific scientific goals.
OBDA2
PEX_2
SYSTEM HOUSEKEEPING
FLIGHT DATA
PAYLOAD
EXPERIMENTS
STRUCTURE
ATD
INSTRUMENTATION TRANSITION BASE FLOW SWBLI S.W. IMPINGMENT
TPS
Fig. 22 – Instrumentation Categories
Fig. 23 illustrates the envisaged architecture showing the main interfaces with the On-Board Data Handling (OBDH) Subsystem.
Experiments Instrumentation is the set of instrumentation distributed over the vehicle configuration to perform investigations over the specific regions. Presently three main experimentation areas can be identified. One belongs to the study of the behavior of the structure during the mission. It consequently foresees a set of sensors
embedded in the structure of the vehicle; the second belongs to experimentation in the filed of Aero-thermodynamics and it is intended to deeply investigate specific aspects such as Transition phenomena, Base drag phenomena, Shock Wave – Boundary Layer Interaction (SWBLI) and Shock Impingement; the third is addressed to the in-flight study of the TPS behavior. As an example, Fig. 24 reports a possible instrumentation distribution of sensors for the analysis of the TPS behaviour over the FTB_2 external surfaces.
Thermocouple (93) Heat Flux (13) Strain Gauge (3)
Fig. 24 – Sensors Location for TPS investigation
Fig. 25 – FTB_1/2 Instrumentation (partial)
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For what concerns the Payload, it has been conceived as a further opportunity offered to the experimenters; such opportunity concretizes as follows: − a volume (schematically drawn as a cube of
0.5[m] per side) of 0.125[m3]; − a mass allocation of 40 [kg]; − a digital channel to the OBDH via 1553 bus; − ten (10) analogic channels to the OBDH via
VME bus. The Payload type of experimental capability can be used to host real independent experiments taking benefit from the environment offered by USV missions. Examples of Payload are: – advanced aerothermodynamic laser-based
diagnostics to deeply investigate the flow field near the vehicle and the effects of surface interaction phenomena (mainly on ORT)
– innovative TPS components (SRT, HFT, ORT) – specific technological elements to be
tested/qualified in flight, such as special wing prototype, propulsion system breadboard (see next chapter)
a microgravity experiment could be executed on SRT tanks to the parabolic-like trajectory eventually optimized to produce enough duration of very low gravity conditions.
ON-BOARD INSTRUMENTATION
OBSW1 is the software component that, within the FTB1 project, shall implement the logical functions required for the configuration and management of the system and of the various subsystems; it shall also incorporate appropriate functionalities to support the exchange of information among the subsystems and between the USV and the EGSE or the Ground Station.
Fig. 26 - OBDH Context Architecture Block Diagram
Fig. 26 shows the general architecture of the system components the OBDH platform shall be interfaced
with. These components, consequently, shall be regarded as the environment that the software application OBDSW1 shall manage. As shown in the Fig. 26 above, the system will communicate with the external world by means of three external data links; in particular while the USV is still on the ground, it will be connected to the Electrical Ground Support Equipment (EGSE) by means of the Umbilical Link; this will be made of: 1. Umbilical Data Link: for the transfer of data
between the OBDH and the GSE and v.v. 2. Umbilical Power Link: for the power supply of
all the USV equipment, through the Electrical Power Supply Subsystem.
While the USV is ascending hanged to the Carrier device, it will be connected to the Carrier Data Handling Unit by means of a serial data link that could be used as alternative data link to the TT&C data link for exchanging messages with the Ground Station.
USV-TECH: THE USV TECHNOLOGY PLAN
Brief highlights on each of the technological projects running under USV-TECH are given hereinafter, together with relevant connections with the FTB missions. SHS: Sharp Hot Structures (former: Advanced TPS with UHTC) – The aim is to design, realize and qualify (both on-ground and in-flight) hot structure components made of innovative ceramics able to operate as advanced TPS at temperatures as high as 2000°C. Preliminary components made of C/SiC, SiC/SiC ZrB2/SiC, oxide based CMC have been produced and are being tested in relevant plasma environment. It is planned to:
fly an on-ground qualified Nose on SRT (2006); fly the already in-flight qualified Nose and on-
ground qualified WLE on HFT (2007) and HFT-LP (2008)
aiming at reaching a Technology Readiness Level (TRL) of 7. CRYOTANK-2: Filament Wound Cryogenic Tank – The project is intended to develop specific know-how to design and build a prototype CFRP tank by means of the Filament Winding process. In particular, it will address mainly the issue of permeability of such composite tanks and, such as that, the attention is focused on liquid hydrogen propellant. Nevertheless, taking into account the interest of the Italian industry, the physical problems associated with LOx tanks will also be addressed. It is matter of fact that the most advanced experimental objective of the project is to
fly a prototype LOx tank on HFT-LP (2008)
trying to arrive at a final TRL = 7. First level design optimization and sample “bottles” were already
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realized and tested under cryogenic inert conditions (Fig. 27).
PROP: Space Propulsion – The development of enabling technologies for RLV main propulsion is the goal of this project. The focus is concentrated on combustion chamber phenomenology of liquid hydrocarbon and solid propellant air-breathing (Scramjet) engine. The connection with the shorter term LOx/Hydrocarbon rocket technology is also maintained and linked to a separate national program funded by the Italian Space Agency (ASI). Target objectives are:
fly the HFT-LP mission (2008) with a 10 t LOx-HC engine to be developed within the mentioned separated program. This test should allow reaching a TRL of 7
fly a breadboard Scramjet experiment on HFT (2007) or HFT-LP (2008) arriving at a final TRL of 5.
ATD: Aerothermodynamics – The discipline is one fundamental brick for the fruitful development of next generation RLVs. It is planned to: (i) develop a new multi-objective design optimization methodology for new aerodynamics configurations; (ii) study of laminar-to-turbulent flow transition to improve the ability to predict this phenomenology; (iii) develop know-how for the Extrapolation-to-Flight problem, with specific regard to shock wave boundary layer interaction (SWBLI). It is expected to
fly laminar-to-turbulent transition experiment/s on SRT (2006), HFT (2007) and/or HFT-LP (2008), ORT (2009)
fly extrapolation-to-flight experiment/s on SRT (2006), HFT (2007), ORT (2009).
I&IHMS: Intelligent & Integrated Health Management Systems – Reusability of future RLVs passes through the development of the capability to continuously monitor and then manage the health status of the system, to avoid dangerous loading conditions and to activate those actions that could prevent any accelerated ageing process of the system. Thus, the project intends to develop and test new diagnostic and prognostic capabilities, as well as
procedures to increase safety, reusability and performances. The target TRL is 5 by
flying an experiment on one of the USV missions. GNC: Autonomous Guidance Navigation & Control – The main aim is here to develop autonomous GNC design capability and systems for either the reentry phase and hypersonic flight. While it is considered not strictly necessary to control the USV reentry missions by means of the developed capabilities (not autonomously controlled missions), it is deemed fundamental implement this technology to fly the horizontal hypersonic flight test/s. Consequently, the target TRL is different for the two applications; thus:
fly a reentry GNC experiment on SRT (2005) and ORT (2008) with a final TRL of 5
fly a hypersonic GNC experiment on HFT (2006) and HFT-LP (2007) with a final TRL of 6.
AHW: Adaptive Hypersonic Wing – Decreasing the dry mass of future RLVs can be contributed by decreasing the mass of single structure components. The project thus aims at studying and developing the enabling technologies for the realization of RLVs “smart” components. In particular, it is intended to design and realize a small hingless hypersonic wing based on shape memory materials able to guarantee assigned shapes in different flight conditions as function of wing temperature. The arrival point is to fly an experiment on HFT (2007) or HFT-LP (2008) reaching thus a TRL of 6
PROGRAM STATUS
After an intense Phase A, the USV program was official started on March 2002. The Phase B of FTB_1 and FTB_2 is actually running since that date, having successfully passed in October 2002 the Preliminary System Requirements Review (P-SRR). The fundamental milestones of the program up to the execution of DTFT are: SRR December 2003 PDR-1 March 2004 CDR-1 June 2004 FTB_0 S/S delivery by September 2004 DTFT_0 AR May 2005 DTFT_0 flight July 2005 AR_1 October 2005 DTFT flight December 2005
AKNOWLEDGEMENTS
The author expresses the acknowledgements to all members of the USV team. In particular: − Pier Paolo De Matteis, Responsible Space
Technologies
Fig. 27 – CFRP Filament Wound “bottle” tank
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American Institute of Aeronautics and Astronautics
− Massimiliano Pastena, Responsible USV_1/ DTFT
− Francesco Curreri, Responsible USV_2/SRT − Michelangelo Serpico, Responsible System
Engineering Section − Valter Pesce, Responsible Project Control
REFERENCE 1) G. Russo et al., “ Preliminary Design and Performance
of the PRORA-USV Experimental Vehicle”, 2nd Int. Symp. on Atmospheric Re-entry Vehicles and Systems, Arcachon 26-29 March, 2001
2) G. Russo, S. Borrelli, G. Borriello, A. Denaro, F. Betti, A. Accettura, “Access to Space: Flying Test Beds as Need for Long Term R&D”, 2nd Int. Symp. on Atmospheric Re-entry Vehicles and Systems, Arcachon 26-29 March, 2001
3) G. Russo, “Next Generations Space Transportation Systems: R&D and Need for Flying Test Beds”, AIAA/NAL/NASDA/ISAS 10th Int. Space Planes and Hypersonic Systems and Technologies Conference, Kyoto (Japan) – 24-27 April 2001
4) G. Russo, “The PRORA-USV Programme”, 4th European Symp. on Aerothermodynamics for Space Vehicles, CIRA, Capua (Italy) 8-11 October 2001, ESA-SP-487, pp. 37-48, March 2002
5) G. Russo, “Next Generations Space Transportation Systems”, presented at MUSEAS1 MUltifunction SEnsors For Structural Health Monitoring In Aerospace Structures, Capua Italy, 8-9 November 2001, Aerotecnica Missili e Spazio, Vol. 81 N. 2, pp. 65-72, April-June 2002.
6) G. Russo, V. Salvatore, “PRORA-USV Space Propulsion Technologies”, 8-IWCP Int. Workshop on Rocket Propulsion: Present and Future, Accademia Aeronautica, Pozzuoli (Italy), 16-20 June 2002
7) G. Russo, “Towards RLVs: the PRORA-USV Program”, 11th AIAA-AAAF Int. Aerospace Plane & Hypersonic Syst. & Techn. Conf., Orléans (France), 29 Sept – 4 Oct 2002
8) G. Marino, G. Russo, A. Denaro, G. Borriello, “USV: Flying Test Bed Opportunity for TPS and Hot Structures”, 4th European Workshop on Hot Structures and Thermal Protection Systems for Space Vehicles, Palermo (Italy), 26-29 November, 2002
9) G. Russo, G. Marino, “The USV Program & UHTC Development”, 4th European Workshop on Hot Structures and Thermal Protection Systems for Space Vehicles, Palermo (Italy), 26-29 November, 2002
10) S. Borrelli, M. Marino, “The Technology Program in Aerothermodynamics for PRORA-USV”, 4th European Symp. on Aerothermodynamics for Space
Vehicles, CIRA, Capua (Italy) 8-11 October 2001, ESA-SP-487, pp. 37-48, March 2002
11) G. Russo, “Status of the PRORA-USV Program”, 3rd Int’l Symposium, Atmospheric Re-entry Vehicles and Systems, Arcachon (France) – 24/27 March 2003
12) F. Curreri, G. Guidotti, G. Russo, A. Sansone, M. Solazzo, “PRORA USV Program – The Suborbital Re-Entry Test”, IAC-03-V.6.06, Int’l Astronautical Congress, Bremen (Germany), 29 Sept.- 3 Oct. 2003
13) F. Curreri, G. Guidotti, G. Russo, A. Sansone, M. Solazzo, “PRORA USV Program – The Suborbital Re-Entry Test”, IAC-03-V.6.06, Int’l Astronautical Congress, Bremen (Germany), 29 Sept.- 3 Oct. 2003
14) G. Marino, D. Tescione et al., “New Materials and Related Fabrication Processes for Hot Structures on RLV’s“, Int’l Astronautical Congress, Bremen (Germany), 29 Sept.- 3 Oct. 2003
15) G. Perrone, R. Sabatano, “FTB_0: The In-Flight Qualification System of the Italian USV Program First Mission “, IAC-03-U.4.05, Int’l Astronautical Congress, Bremen (Germany), 29 Sept.- 3 Oct. 2003
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