IRENE PROGRAM

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Aerospace Laboratory for Innovative components IRENE PROGRAM I I talian talian R R e- e- E E ntry ntry N N acell acell E E Preliminary Study Preliminary Study 1 1 0th INTERNATIONAL Planetary Probe WORKSHOP 0th INTERNATIONAL Planetary Probe WORKSHOP June 17-21, 2013 June 17-21, 2013
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IRENE PROGRAM. I talian R e- E ntry N acell E Preliminary Study. 1 0th INTERNATIONAL Planetary Probe WORKSHOP June 17-21, 2013. BACKGROUND OF THE ACTIVITY (1/2). - PowerPoint PPT Presentation

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PowerPoint PresentationIRENE PROGRAM
10th INTERNATIONAL Planetary Probe WORKSHOP June 17-21, 2013
Aerospace Laboratory for Innovative components
BACKGROUND OF THE ACTIVITY (1/2)
The Italian Space Agency (ASI) is supporting since 2010 a research programme, called IRENE, in Campania (ALI, South of Italy), to develop a low-cost re-entry capsule, able to return payloads from the ISS to Earth and/or to perform short-duration, scientific missions in Low Earth Orbit (LEO).
The main features of the IRENE capsule are:
• light weight (100-200 kg), 3 m fully deployed
• payload recoverability and reusability
a fixed nose (made by ceramic or other equivalent TPS)
a deployable aero-brake (umbrella-like, made by special multi-layered fabric).
ALI - Aerospace Laboratory for Innovative components is as a Consortium of17 Companies operating within the fields of design, engineering, prototyping and realization of innovative aerospace sybsystems and Ground Segment for technological and scientific platforms
Aerospace Laboratory for Innovative components
BACKGROUND OF THE ACTIVITY (2/2)
The feasibility study of this deployable re-entry system has been carried out in 2011.
The TPS materials, selected for the nose cone and for the flexible umbrella shield, have preliminarily been tested in the SPES hypersonic wind tunnel at the University of Naples, and in the SCIROCCO PWT (Plasma Wind Tunnel) at CIRA (Centro Italiano Ricerche Aerospaziali) of Capua, Italy.
IRENE TPS test in the SCIROCCO Plasma Wind Tunnel at CIRA
Aerospace Laboratory for Innovative components
MINI-IRENE DEMONSTRATOR (1/2)
On the basis of the previous results, ESA supported a six months "Bridging Phase” to preliminarily address the main issues of a MINI-IRENE demonstrator to be embarked as a piggy-back payload for a future mission of a sub-orbital MAXUS sounding rocket.
The Mini-IRENE system shall be boarded as a secondary payload in the inter-stage adapter of the rocket and ejected, at an altitude of about 150 km, to perform 15 minutes ballistic flight.
A possible launch of a demonstrator of IRENE from a sounding rocket will require scaling down the most important parameters
Aerospace Laboratory for Innovative components
MINI-IRENE DEMONSTRATOR (2/2)
Considering a cylindrical volume available with D=29 cm, h=25 cm, mass≈15-20 kg, the following issues have been addressed:
Analysis of the time profiles of the different physical parameters of interest (e.g. pressure, temperature, acceleration).
Preliminary aerodynamic and aero-thermodynamic analysis (engineering methods and CFD analysis of 45 and 60 deg half cone)
Identification of the main mission requirements and corresponding subsystems
Trade-off between different configurations and identification of possible solutions for the different subsystems.
Aerospace Laboratory for Innovative components
AEROTHERMODYNAMIC ANALYSIS
Pressure distributions at maximum dynamic pressure condition (left: maximum 21 kPa, right: maximum 10.5 kPa)
Main geometric and aerodynamic characteristics
same length of the poles
Half-cone angle [°]
D [m]
S [m2]
MINI-IRENE REQUIREMENTS
Maximum diameters of 29cm (folded) 100cm (deployed)
Total mass below 20 kg / Ballistic coefficient less than or equal to 18 kg/m
Deployable heat shield
Automatic system for TPS deployment (during exo-atmospheric phase) generating a 45-60 deg sphere-cone shape
Structure able to withstand mechanical loads at launch and aerodynamic loads during reentry (10000 Pa stagnation pressure, 40g deceleration, impact loads at landing with a velocity in the order of 20 m/s)
TPS able to withstand heat fluxes in the order of 350 kW/m2
CoG location to guarantee stability and reduce trim angle
Aerospace Laboratory for Innovative components
Three possible solutions have been considered for the supporting structure:
Telescopic poles
Folded arms
Hinged arcs
Any of the solutions foresees upper and lower threads to give rigidity to the whole system and a mobile structure (≈ Tensegrity)
Sufficient room, within the dedicated Maxus volume, is left for the TPS fabrics that must stay properly bended before the deployment
In the present preliminary design all constructions includes a series of 12 main elements (poles, arms or arcs) but this number may vary
PRELIMINARY DESIGN OF THE SUPPORTING STRUCTURE
Aerospace Laboratory for Innovative components
Solution 1: 12 telescopic poles
Closed
a1) folded structure; b1) pole elongation phase; c1) tensioning phase
Poles hinged to sliding structure
Upper threads anchored to fixed structure
Lower threads anchored to sliding structure
Elongated
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a2) folded structure; b2) arm extension phase; c2) tensioning phase
Two-segment
Upper threads anchored to fixed structure
Lower threads anchored to sliding structure
Aerospace Laboratory for Innovative components
Solution 3: hinged arcs
Bended arcs of ring
a3) folded structure; b3) rim extension phase; c3) tensioning phase
Example of double universal joints. Similar joints could be used to connect the elastic arcs of the ring
Arcs hinged one another
Aerospace Laboratory for Innovative components
TPS PRELIMINARY DESIGN
- a rigid nose
- a flexible part, to be deployed prior to re-entry.
The flexible part of the heat shield is not requested to function much as a thermal insulator but, rather, mainly as an aero-brake.
The deployable part of the thermal shield should be sufficiently thin and flexible for an easy deployment.
It will be necessary to prove that the exposure of the proposed materials to the typical heat fluxes expected during descent would not compromise the tensile strength of the flexible part of the heat shield.
Aerospace Laboratory for Innovative components
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PRELIMINARY DESIGN OF DEPLOYMENT MECHANISM (1/2)
Two different mechanisms have been considered to deploy the IRENE aerobrake/heat shield:
Mechanism #1: Including actuator springs and gas dampers
Mechanism #2: Including a gas actuator
Description of Deployment Mechanism #1
This solution exploits harmonic steel compression springs that, once loaded, store the needed mechanical energy to perform the heat shield deployment. In order to avoid abrupt elongation of springs as they are unlocked, a damping system has been devised.
Description of Deployment Mechanism #2
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PRELIMINARY DESIGN OF DEPLOYMENT MECHANISM (2/2)
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Aerospace Laboratory for Innovative components
A FEM structural analyses has been performed for each of the three identified solutions. The main goal of the activities is the evaluation of the stability and the stress levels in the most critical components of the three solutions. The components so identified, are as follows:
Poles (sol. #1), arms (Sol. #2) and arcs (Sol. #3);
Threads
TPS Fabric layers (FEM models, four layers of NEXTEL AF-10, thickness=0.39 mm).
Results: admissible stress levels for the structural components (considering also the operating temperature) is the following:
Titanium structure 400 MPa at 400°C
NEXTEL Fabric : 40 MPa at 900 °C
STRUCTURAL ANALYSIS
Aerospace Laboratory for Innovative components
The solution #1 (with 45 deg half cone) shows a better behavior for the following aspects:
SOLUTION SELECTION
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
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INSTRUMENTATION
The most important parameters to be quantified during the re-entry are the aerothermodynamic loads, represented by the surface pressure distribution and the surface heat flux.
The main assumption is that the payload consists only of the following sensing elements and their respective data acquisition and storage electronics and power supply:
• Thermocouples
• Strain gages
• Telecommunications subsystem
• On Board Data Handling
Sensors: thermocouples location
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.
INSTRUMENTATION: THERMOCOUPLES
Payload temp.
Computer temp.
Cone shape
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.
INSTRUMENTATION: PRESSURE SENSORS
AVIONICS AND INSTRUMENTATION
Telecommunication / Data Retrieval
No TMTC. Recovery of the capsule via beacon (to be choosen/developed/upgraded).
GNC
Trajectory measurement by use of MEMS based IMU and eventual additional axial accelerometer for high accelerations.
Data Handling
Data concept
Image and sound recording
The presence on board of a video camera and a microphone is useful as additional check on the functioning of the experiment.
Power
No telemetry
A beacon system will be used for the recovery after landing
The beacon shall be operational before landing
The beacon shall be operational after a landing for at least 48hours
The use of standard call and Search and Rescue system, such as Cospas-Sarsat, is allowed
The baseline configuration is “integrated Beacon with antenna out of back TPS”
Trade-off compared and help to select COTS beacon.
Main issues are:
Small and light equipment.
Crushability requirements are critical, in order to avoid previous “Shark mission” impact problems
Power autonomy for at least 48hours
TELECOMUNICATION
Aerospace Laboratory for Innovative components
Trade-off compared and help to select the GNC equipment that better match mission requirements. Two different options:
Individual accelerometers and gyroscopes selected independently (no IMU) (best solution: solution n. 2 is not readily usable because of the difficulty of interfacing with the OBDH system selected).
Integrated IMU with an acceleration range of more than 40g
To acquiring flight parameters and reconstructing flight trajectory of the capsule
Acceleration requirement of a minimum of 40g, (maximum re-entry Acceleration)
Detection of linear and rotation rate of acceleration
Minimize mass and weight
Aerospace Laboratory for Innovative components
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 crash-landing.
Two design possibilities have been evaluated:
Embed the VMU in the Data Handling System. DHS can provide sufficient memory to allocate 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, the first configuration was selected (flight proven, ACRA KAM 500 modular computer, able to acquire and store, on a ruggedized memory unit, all the data)
2) Other projects (“Phoebus”) in order to be more resistant to crash, decided to implement the VMU as a separate unit interfacing the DHS using USB bus.
ON BOARD DATA HANDLING
Preliminary functional diagram

MAXUS-8 Switch Interface
PRELIMINARY TOTAL MASS BUDGET
Future Development
About development plan for the MINI IRENE project up to launch, presently planned in the first half of 2015, is detailed below. It includes the already performed Preliminary Study (referred to as “Bridging Phase”, BP, in the figure below) which concludes the Phase 0 studies previously performed on behalf of ASI.
Payload temp.
Computer temp.
Cone shape