IRENE PROGRAM. European Sounding Rocket Experiment on Hypersonic Deployable Re-entry Demonstrator

Transcription

1 IRENE PROGRAM European Sounding Rocket Experiment on Hypersonic Deployable Re-entry Demonstrator R. Savino, R. Aurigemma, Dr. Pasquale Dell Aversana, L. Gramiccia, F. Punzo, J. Longo, L. Scolamiero, L. Marraffa 8th European Symposium on Aerothermodynamics for Space Vehicles 2-6 March 2015, IST Lisbon, Portugal 1

2 BACKGROUND OF THE ACTIVITY (1/2) ASI and ESA are supporting since 2010 IRENE (Italian REentry Nacelle for Experiments), to develop in Campania (South of Italy) a low-cost re-entry capsule, able to return payloads from the ISS to Earth and/or to perform short-duration, scientific missions in LEO. IRENE capsule main features : light weight ( kg), 3 m fully deployed payload recoverability and reusability low-cost deployable disposable heat shield composed of: o a fixed nose (special ceramic material) o a deployable aero-brake (umbrella-like, special multi-layered fabric). ALI - for Innovative components is as a Consortium of 17 Companies operating within the fields of design, engineering, prototyping and realization of innovative aerospace sybsystems and Ground Segment for technological and scientific platforms 2

3 BACKGROUND OF THE ACTIVITY (2/2) Feasibility study carried out in TPS materials, selected for the nose cone and for the flexible umbrella shield, preliminarily tested in SPES hypersonic wind tunnel (University of Naples), and in SCIROCCO PWT at CIRA (Centro Italiano Ricerche Aerospaziali) of Capua, Italy. IRENE TPS test in the SCIROCCO Plasma Wind Tunnel at CIRA 3

4 MINI-IRENE DEMONSTRATOR (1/2) Based on previous results, ESA supported a 6-months "Bridging Phase : preliminarily design of MINI-IRENE demonstrator to be embarked as a piggy-back payload of a sub-orbital MAXUS sounding rocket. Mini-IRENE system: secondary payload in the inter-stage adapter of the rocket ejected, at an altitude of about 150 km, to perform 15 minutes ballistic flight. Launch of IRENE demonstrator from a sounding rocket requires scaling down most important parameters 4

5 MINI-IRENE DEMONSTRATOR (2/2) Available volume : cylinder D=29 cm, h=25 cm, mass kg, Issues addressed: Analysis of time profiles of relevant parameters (e.g. pressure, temperature, acceleration). Preliminary aerodynamic and aerothermodynamic analysis (engineering methods & CFD for 45 and 60 half cone) Identification of main mission requirements and corresponding subsystems Trade-off between different configurations and identification of possible solutions for the different subsystems. 5

6 AEROTHERMODYNAMIC ANALYSIS Altitude versus time Altitude versus acceleration Altitude versus stagnation pressure Altitude versus stagnation point convective heat flux. Half-cone angle [ ] D [m] S [m 2 ] c D = m/(c D *S) Pressure contours at max. Pdyn. conditions (left: 45 half-cone angle, Pmax 12 kpa; right: 60 half-cone angle, Pmax 8 kpa) Main geometric and aerodynamic characteristics 6

7 MINI-IRENE REQUIREMENTS Compatibility with MAXUS inter-stage Ø max = 29cm (folded) 100cm (deployed) Total mass 15 kg / Ballistic coef. β 20 kg/m 2 Deployable heat shield Automatic system for TPS deployment during exo-atmospheric phase generating a sphere-cone shape Structure: Mechanical loads at launch and aerodynamic loads during reentry (12 kpa stagnation pressure, 35g deceleration, impact loads at landing with a velocity in the order of 20 m/s) TPS: heat fluxes in the order of kw/m 2 CoG location to guarantee stability and reduce trim angle 7

8 PRELIMINARY DESIGN OF THE SUPPORTING STRUCTURE Different solutions considered for TPS supporting structure. Two basic concepts in common: The deployable part of the heat shield is a tensile structure; Deployment is performed in two phases: coarse elongation, characterized by long strokes (of poles or arms) and relatively weak forces; tensioning, characterized by shorter strokes (of a sliding structure) and stronger forces. 8

9 Solution selected: 12 telescopic poles Poles hinged to sliding structure Upper threads anchored to fixed structure Lower threads anchored to sliding structure Upper threads Sliding structure Nose (ceramic) TPS Fixed structure Multi layer TPS fabric Closed telescopic poles Elongated telescopic poles Lower threads a 1 ) folded structure; b 1 ) pole elongation phase; c 1 ) tensioning phase 9

10 TPS PRELIMINARY DESIGN The TPS composed of two main sections: -a rigid nose - a flexible part, deployed prior to re-entry. Flexible part of heat shield: aero-brake more than thermal insulator. Deployable part of thermal shield should be sufficiently thin and flexible for an easy deployment. Necessary to prove that exposure of proposed materials to typical entry heat fluxes does not compromise tensile strength of flexible part of TPS. 10

11 MINI IRENE LAYOUT (1/3) Mine-Irene 3D view in stowed configuration 11

12 MINI IRENE LAYOUT (2/3) Mini-Irene 3D section view in extended configuration 12

13 MINI IRENE LAYOUT (3/3) Mini-Irene 3D section view in opened configuration 13

14 Solution selected: 12 telescopic poles VIDEO\Apertura fase 2.avi 14

15 Solution selected: 12 telescopic poles VIDEO\Apertura fase 2.avi 15

16 STRUCTURAL ANALYSIS A FEM structural analyses has been performed. Main goal : evaluation of structural stability and of stress levels for most critical components of selected solution. Identified components : Poles Threads TPS Fabric layers (FEM models, four layers of NEXTEL AF-10, thickness=0.39 mm). Results: admissible stress levels for structural components (considering also operating temperature) : Titanium structure 400 MPa at 400 C NEXTEL Fabric : 40 MPa at 900 C 16

17 SOLUTION SELECTION Advantages of selected solution : Better aerodynamic stability, due to smaller cone angle (45 instead of 60 ) Largest diameter of the deployed structure due to the deploying mechanism kinematics Better fabric tension distribution after the deployment phase due to deploying mechanism kinematics Lower fabric deflection under the re-entry pressure loads 17

18 INSTRUMENTATION Most important parameters to be quantified during the re-entry: aerothermodynamic loads, i.e. surface pressure distribution and surface heat flux. Payload consists of the following sensing elements and their respective data acquisition, handling & storage electronics and power supply: Thermocouples Pressure transducers accelerometers and gyroscopes Strain gauges Telecommunications subsystem On Board Camera and Sound Recorder 18

19 INSTRUMENTATION: THERMOCOUPLES Sensors: thermocouples location 2-3 thermocouples at the stagnation point at different depth 2 thermocouples embedded in outer positions of the nose cone 3 thermocouples at different positions of the flexible shield, (embedded in the last inner textile layers). 3 thermocouples located out of the capsule body, and 2 inside the payload area 3 thermocouple will monitor the ribs heating. 19

20 INSTRUMENTATION: PRESSURE SENSORS Sensors: pressure sensors location 3 pressure sensors embedded at the stagnation point and in (2) outer positions of the nose cone. 2 pressure sensors located at different positions out of the capsule body, inside the cone area. 2 pressure sensors located on the back shield of the capsule body. 20

21 AVIONICS AND INSTRUMENTATION Telecommunication / Data Retrieval GNC Data Handling Data concept Image and sound recording Power No TMTC. Recovery of the capsule via beacon (to be choosen/developed/upgraded). Trajectory measurement : MEMS based IMU and eventual additional axial accelerometer for high accelerations. Preferably COTS electronics. Possibly merged with power conditioning / distribution to provide miniaturization. 1 Processor boards (OBC, Payload) 2 Analog acquisition board, 1 output board Recoverable on board recording, no Telemetry video camera and a microphone as additional check on the functioning of the experiment. Primary batteries only 21

22 AVIONICS AND INSTRUMENTATION Preliminary functional diagram 22

23 PRELIMINARY DEFINITION OF AVIONICS AND INSTRUMENTATION Mini-Irene Avionic systems section view 23

24 TELECOMMUNICATION No telemetry Beacon system for recovery after landing: The beacon shall be operational before landing The beacon shall be operational after landing for at least 48hours Standard call and Search and Rescue system, such as Cospas-Sarsat, is allowed Baseline configuration: integrated Beacon with antenna out of back TPS COTS beacon selection main criteria: Small and light equipment. Crushability requirements are critical, to avoid previous Shark mission impact problems Power autonomy for at least 48hours 24

25 GNC Acquisition of flight parameters for capsule flight trajectory reconstruction Acceleration requirement of a minimum of 40g, (maximum re-entry Acceleration) Detection of linear and rotation rate of acceleration Minimize mass and weight Trade-off to select the GNC equipment. Two different options: Individual accelerometers and gyroscopes selected independently (no IMU) (best solution) Integrated IMU with an acceleration range of more than 40g, (difficult interfacing with the OBDH system selected). 25

26 ON BOARD DATA HANDLING The Vehicle Memory Unit (VMU) is one of the most critical part of the mission: it will store all data and has to survive the crashlanding. Two design possibilities have been evaluated: Embed the VMU in the Data Handling System. DHS provides sufficient memory to store the whole mission data. The whole DHS have to survive the crash-landing. Consider the VMU as a separate unit with an external interface to the DHS. 1) In previous Shark mission, first configuration selected (flight proven, ACRA KAM 500 modular computer, to acquire and store all the data on ruggedized memory unit,) 2) To be more resistant to crash, other projects ( Phoebus ) implement VMU as separate unit interfacing DHS using USB bus. 26

27 GROUND DEMONSTRATOR A technology ground demonstrator of the umbrella-like structure developed to test on ground most critical functions (kinematics of the deployment system, configuration of the flexible structure, efficiency of the selected actuators, etc.) and the mechanical stresses expected in flight. Mini-Irene Ground Demonstrator 27

28 GROUND DEMONSTRATOR: TEST The following tests have been performed: Permeability Test Dimensional Test Telescopic Pole Elongation Test Functionality Test of Deployment System Release Mechanism Test Pressure Loads Test Vibration Test The demonstrator at maximum pressure loading (6,4 kpa) in test session 28

29 TOTAL MASS BUDGET 29

30 Future Development Development plan for MINI IRENE up to launch, presently planned in the first half of 2017: 30

31 Conclusion IRENE techno ground demonstrator successfully tested Next step: flight demonstration A number of potential applications identified. First one, LEO re-entry 31

32 Thank you! 32