recent advancements in the development and … · recent advancements in the development and...

Recent Advancements in the Development and Verification of

Space Vehicle Structures

Prof. Dr. C. StavrinidisHead of Mechanical Engineering Department

([email protected])

European Space Agency

Future Trends in Certification of Advanced Technology Structures, Bristol, 16 Sep 2015

Presentation Objectives

1. Overview of Mechanical System oriented developments at ESAa. The Concurrent Design Facility (CDF)

b. Spacecraft mechanical environment and test verification

c. Vibro-acoustic Evaluation

d. Shock environment evaluation

e. Herschel telescope thermo-elastic evaluation

f. Satellite High Pointing Stability Requirements

2. Overview of current/recent composites structures technology developments in ESA for spacecraft and launcher structures applications and related Verification/Certification challenges:

a. Design & Verification

b. Composites Applications for Spacecraft

c. Composites Applications for Launchers

d. Composite technology developments

e. Analysis tools and methodologies

Future Trends in Certification of Advanced Technology Structures, Bristol 16 September 2015

“To provide for and promote, for exclusively peaceful purposes, cooperation among European states in

space research and technology and their space applications.”

Article 2 of ESA Convention

PURPOSE OF ESA


ESA FACTS AND FIGURES

• Over 50 years of experience

• 20 Member States

• Eight sites/facilities in Europe, about 2200 staff

• 4.4 billion Euro budget (2015)

• Over 80 satellites designed, tested and operated in flight

• Over 20 scientific satellites in operation

• Six types of launcher developed

• 200th launch of Ariane celebrated in February 2011



20 MEMBER STATES AND GROWING

ESA has 20 Member States: 18 states of the EU (AT, BE, CZ, DE, DK, ES, FI, FR, IT, GR, IE, LU, NL, PT, PL, RO, SE, UK) plus Norway and Switzerland. Estonia and Hungary will soon be part of ESA (2015)

Seven other EU states have Cooperation Agreements with ESA: Bulgaria, Cyprus, Latvia, Lithuania, Malta, Slovakia and Slovenia. Discussions are ongoing with Croatia.

Canada takes part in some programmes under a long-standing Cooperation Agreement.


Mechanical System oriented developments:The Concurrent Design Facility (CDF)Where it all begins……

The size and shape of the spacecraft configuration is dictated by a large number of requirements addressing:

a. Spacecraft mission (orbit, propulsion system and size, power requirements):- how many tanks and their size; - solar array size;

b. Payloads and equipment (size, pointing, environmental requirements):-size of payloads;- location;- thermo-elastic stability requirements;

c. Launcher (interface, loads, volume):- the physical interface to the launcher;- the stiffness requirement imposed by the launcher

(no dynamic coupling)- space available under the fairing

Initial design concept is facilitated through the use of Concurrent Design Facility (CDF)


Mechanical system oriented developmentThe Concurrent Design Facility (CDF)

• The ESTEC Concurrent Design Facility is an Integrated Design Environment available to all ESA programmes for interdisciplinary and inter-directorate applications, based on a Concurrent Engineering approach!

• Featuring:• concurrent engineering;• integration of tools and project data;• mission and system models;• simultaneous participation of all

subsystem and system disciplines;

• Scope now extended to include payload instrument conceptual design, scientific requirements definition and consolidation, reviews of industrial phase A studies,anomaly investigation as well as for education and training.


Human Missions to Mars

MoonLander

ROSITA instrument on Columbus External Platform

ISS Internal Payload –Science Requirements Definition

IMPACT facility inside an ISS rack

Socrates

Heavy Lift Launch Vehicle

Crewed vehiclesfor exploration

preparation programme

Telescopes and Technology

FIRI

WiFLY

Diverse range of space missions

S5P

ExoMars

Laplace

Payload &P/L accommodation

Advanced launchers

Mechanical system oriented developmentsCDF studies and achievements


Mechanical System oriented developments: Spacecraft mechanical environment and test verification 1/7

Launch loads:• Quasi static acceleration loads (< 10g)• Dynamic loads (+/- a few g’s base input)

(0- 100 Hz)• Random and Acoustic loads

(up to app 1 g2/Hz , app 140 dB SPL)(0 – 2000Hz)

• Shock loads(several thousand g’s but for a short period, few ms)(up to 10000Hz)

For design, the combined static load can be several tens of g’s

On-orbit loads:• Thermal loads:

(from cryogenic temperatures to app 120C)• Attitude Control Loads (mainly relevant for appendages such • as antennae, solar arrays etc);• Meteroid and Debris impact loads; (integrity and disturbance)• Micro-vibration disturbances from moving equipment, e.g. reaction

wheels!

Re-entry / Landing loads:• Pressure loads• Thermal loads (up to app 1700C for exposed hot structures)• Landing impact (earth or planetary)

Subsystem mechanical environment derived from past experience OR by Simulation !


Thermal, vacuum, sun, no-gravity, radiation

Vibration-shock

Noise-vibration

Mechanical system oriented developmentsCDF studies and achievements



BepiColombo – ESA’s mission to Mercury

• ESA mission in cooperation with the Japanese Space Agency – JAXA, will explore Mercury, the closest planet to the Sun.

• Mission will comprise two spacecraft: ESA’s Mercury Planetary Orbiter (MPO) & JAXA’s Mercury Magnetosphere Orbiter (MMO).

• BepiColombo will be launched in 2017 on Ariane 5.

• Spacecraft faces challenging thermal environment at Mercury – 10 times solar radiation at Earth orbit, leading to local temperatures >400°C.

• Large delta-V to overcome Sun’s gravity and enter Mercury orbit requires multiple planetary fly-bys and solar electric propulsion provided by dedicated Mercury Transfer Module (MTM).

• Spacecraft launch mass – 4100kg (includes 1400kg of propellant) • BepiColombo spacecraft Structure/Thermal

Test Model (STM) and Flight Model (FM) are mechanically and thermally tested at ESTEC test facilities.

• STM mechanical testing has been successfully completed in 2012.

• FM spacecraft testing will take place in 2016.



BepiColombo – STM Mechanical Testing

• STM spacecraft Mechanical Qualification programmecompleted in 2012 at ESTEC test facilities.

• Sine vibration tests in three axes from 5-100Hz using Quad and Multi-shaker.

• Acoustic test in LEAF chamber to Ariane 5 acoustic qualification spectrum.

• SHOGUN shock test to simulate Ariane 5 fairing separation shock.

• Clampband firing and Launch Vehicle Adapter (LVA) separation test.

• Module separation tests – firing release mechanisms.

• Mechanical test instrumentation:

- 464 accelerometers, including 112 rated for moderate to high shock inputs.

- 70 strain gauge channels.

- Force measurement device (comprising 24 load cells) at spacecraft base to measure and control base bending moment and force in sine vibration tests.

BepiColombo STM Spacecraft on ESTEC QUAD Shaker



BepiColombo – Thermal Testing

• STM spacecraft modules have completed thermal qualification test programme during 2011-2012 at ESTEC test facilities.

• Modules are tested in ESTEC LSS (Large Space Simulator). Sun simulator beam is reshaped to reproduce sun intensity experienced at Mercury.

• Measured heat flux in test is 13 kW/m2 (10 times level at Earth orit from Sun).

• Testing confirms capability of thermal control systems to maintain internal temperatures within operational limits.

• FM spacecraft modules currently undergoing thermal testing:- MPO FM module completed testing end-2014

(test duration in chamber 19 days).

- MTM FM module will enter testing end-2015.

a. Thermal balance (for MPO module) test instrumentation:

- 584 thermocouples.

- 40 heater lines to simulate internal heat dissipation.

BepiColombo STM MPO 10 Solar Constants Thermal Test in LSS

BepiColombo STM MMO (on spin table)10 Solar Constants Thermal Test in LSS



Dynamic simulation on a portion of an AIRBUS fuselage in a frame of a study carried out by Airbus to investigate structure responses to high frequency excitation : 6-DOF transient simulation.



• Measure the interface forces during base excitation vibration tests.

• Provide the 6 main components of the force/moment vector at the interface point.

• Use the global force/moment signals in real time to control or limit the vibration excitation level.

• Up to 24 individual load-cells can be combined to support various standard interfaces as well as customized setups.

Force Measurement Device FMD



Mechanical system oriented developmentsVibro-acoustic Evaluation 1/3

Equipment loads specifications have to be derived in the early design process with very limited information on the actual configuration of the spacecraft. Specifications need to be adequate but not too conservative.

Early definition of the specification based often on heritage of test data from other spacecraft and on empirical data [ECSS-E-10-03A]

In some situations (e.g. spacecraft of unique dimension or shape), neither the heritage nor the empirical data are considered adequate ! A vibro-acoustic analysis is then performed.

Vibro-acoustic analysis is also performed in order to re-evaluate the equipments’ specifications or to perform trouble shooting in case of anomaly.

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10 100 1000 10000Frequency, Hz

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Typical Acoustic Qualification SPL (141db)

Empirical formula from ECSS

SWARM S/C and its Optical Bench

Optical filter in Proba 2 and predicted pressure level

Vibro-acoustic environment of spacecraft


Mechanical system oriented developmentsVibro-acoustic Evaluation 2/3

Coupled BEM/FEM Analysis ProcedureTypical starting point is the FEM and its modal base.

BEM created as separate entity. Size of BE respects the maximum 1/6 criterion.

Frequency independent viscous modal damping or frequency dependent equivalent viscous modal damping can be considered.(useful in the presence of low damping materials such as SIC and Zerodur)

Diffuse field excitation is represented by uniformly spaced plane waves

Coupled system solved in Windows or Linux environment.

Response spectra/contour in terms of pressure, acceleration, force, or stresses.

Viscous Damping Profiles

0.00

0.50

1.00

1.50

2.00

0.00 100.00 200.00 300.00 400.00 500.00 600.00 700.00Frequency, Hz

%

Case A - Constant Viscous Damping, %Case B - Equivalent Viscous Damping, %

FEM + eigenvectors BEM Fluid Surfaces Modal Damping Profile of Structure

User Inputs

Process Definition GUI (E.g. Diffuse Field, Resolution Steps, Coupling, etc.)

Fluid and SPL Parameters

High Frequency SEA Solver

Low frequency Coupled BEM/FEM

Solver

Mid-frequency Coupled Solver

Results

optional optional


Mechanical system oriented developmentsVibro-acoustic Analysis 3/3

Application Example : HerschelThe HERSCHEL spacecraft weighs about 3250 kg at launch and it is about 7 meters in height. The spacecraft carries a 3.5 m aperture telescope and three scientific payloads operating at near absolute zero K. The payloads (HIFI, PACS & SPIRE) themselves are located on the optical bench which is encapsulated by a vacuum cryo-vessel.FEM has over 1.8 million dofs and high modal density

Test in Blue ; Analysis in Red


Mechanical system oriented developmentsShock environment evaluation

Adequate guidelines for spacecraft and equipment shock analysis are provided in ECSS Mechanical Shock Design and Verification Handbook, ECSS-E-HB-32-25A, covering:

Analytical methods based on attenuation rules, as function of the excitation types (point source, clampband, LV induced shocks). All proposed methods benefit from numerous and rigorous testing on ESA satellites.

Numerical methods (FE or SEA based) for all types of shock excitations, and including shock CLA - Relevant in situations where the shock environment is expected to be influenced by the response of main global modes.

Case of shock responses in large instrument, for which the responses are driven by its dynamic behavior (2 methods are defined, transmissibility and transient analysis approaches).

Equipment shock analysis – In support to design definition and in support to anomaly justification, shock analysis at equipment level is proven to have a significant added value. Such analysis requires that a sufficient correlation is achieved.

Derivation of a consolidated shock environment is an achievable goal, under the condition that a consistent approach is followed.

For more information, See ECSS Mechanical Shock Design and Verification Handbook.


Mechanical system oriented developments:Herschel telescope thermo-elastic evaluation

Large initial discrepancy between the prediction of the change in back focal length of the Herschel telescope and the results obtained from the cryo-optical test!

Tiger Team setup to determine cause of the discrepancy (metrology 11.7 mm, analysis 1.7 mm)

Conclusions: Metrology OK, Models generally OK, but Materials coefficient data not sufficiently accurately measured. Distortion is dominated by the difference in CTE between Silicon Carbide (SiC) and INVAR (ΔCTE)

Flight solution implemented: Shim the telescope (based on metrology data) to place the cryo-focus within the uncertainty volumes of the instruments focal positions

Spacecraft Launched,operating successfully!

ADEQUATE NON LINEAR DEFINITION OF THE COEFFICIENT OF THERMAL EXPANSION (CTE) OF TELESCOPE MATERIAL IS ESSENTIAL FOR SYSTEM VERIFICATION

INVAR STRAIN-TO-CRYO


# Hipparcos 0.300 arcsec 1 minute

# Euclid 0.027 arcsec 12 minutes

# Hubble 0.007 arcsec 18minutes

Satellite High Pointing Stability Requirements

=>Classical “aerospace” type structural verification.

• Advanced numerical modeling

• Supporting subscale development tests

• Ultimate strength safety factors

• Traditional failure models

• Post buckling

• Damage tolerance verification

Future trends:

• Integrated system loads iterations (launcher CLA);

• Non-linear stability (buckling) verification;

• Probabilistic strength verification;

Composites Applications in Spacecraft & Launchers:Design & verification


Why use Composite structures ?

Composite structures offer several advantages depending on the application:

– Mass (spacecraft and launchers);

– Stiffness (mainly spacecraft);

– Strength (mainly launchers);

– thermo-elastic stability (spacecraft);

Composite structures have become an integral part in the implementation of space vehicle developments!


Composites Applications for Spacecraft

Primary / secondary structures

Solar array substrate Antenna reflectors

Equipment

Fasteners & inserts

icotec ag, Switzerland

Airbus DS Spain

ESA/ESTEC The Netherlands


Composites Applications for Spacecraft: Satellite main Structure component

Example: Lisa PathFinder

• Propulsion module central cone (right):

• CFRP filament wound sandwich with Al honeycomb core

• Science module (left):

• Central cylinder, (shear) panels, solar panel

• CFRP face sheet, Al honeycomb core

Launch scheduled for later this year


Example of spacecraft primary structure

Mecabus / Spacebus 4000 Central Tube

(RUAG Sweden & TAS France)

•The CFRP shell is a co-cured sandwich, with internal and external M18/M55J carbon skins, with thickness varying between 0.3 mm and 4 mm, and with various core densities and thicknesses.

•Length of app 4m

•Local skin reinforcements are implemented, either co-cured with the blank shell or cold-bonded on the skins.

•Hundreds of inserts are implemented for accommodation of equipment such as tanks, platforms, flanges and tubes.

•Manufacturing process: Hand lay-up.

•Generally metallic interface rings are used, however composite options exist, e.g. RTM manufactured CFRP rings by CASA


Composites Applications for Spacecraft: Solar Panels, Antennae

Solar Array substrates:

• Aluminum honeycomb core

• High modulus CFRP face sheets

Antennae:• Composites constitutes the basic building block for

advanced antenna reflectors.

Materials: • UHM carbon fibers (M55J, M60J, YSH50…) and new

generation cyanate ester resin systems (M18, 950-1, LY556, EX1515, RS3…) -> microcracking, moisture absorption, thermal and thermo-elastic behaviour, stiffness.

• Prepregs of UD, biaxial fabric or triaxial open weave.• In case of RF transparency Kevlar or Quartz fibres

typically used.

• Pitch Carbon fibres provide an improved thermal and electrical conductivity, as well as elastic modulus. An example of pitch Carbon fibres with cyanate ester resin implementation is the Highly Stable Antenna Surfaces developed under ESA contract(HPS and Invent GmbH Germany).

HSASS (HPS) antenna reflector in DLR TV chamber for distortion measurement with Laser Speckle interferometer.


Antenna Structures

3.8-m shaped reflector antenna (ASTRIUM-ST+TAS-F)

2.3 m ASTRA 1 M shaped reflector (EADS-CASA (E)

2.1 m Shaped dual gridded reflector for C-and Ku-bands (ASTRIUM-ST)

Modular ULR for Eurostar 3000 (ASTRIUM-ST)

Composite reflectors

EXPRESS MD1&2 1,5 m Deployable antenna (Thales Alenia Space Italy)


Composites Applications in Launchers:Ariane 5 Composites Parts

Intermediate cone (only

Ariane 5 ECA)(Airbus DS Spain)

PAS cone (Payload

attachment structure)

Inter Stage Skirt(Airbus DS Spain)

Front skirt (only inner skin part)

GAT/GAM High pressure vessels

EPS structure (mechanical structure & 2 Helium tanks)

(only Ariane 5 ES/ATV)

Ariane 5 ECA

Payload Fairing

VEB (vehicle equipment bay)

Sylda (for dual P/L)

(MT Aerospace)

(Airbus DS Spain)

(Airbus DS)

Thermal protection cover of Vulcain engine

(MT Aerospace)

(MT Aerospace)

(RUAG)

(Airbus DS Spain /RUAG)

(Airbus DS)


Composites Applications in Launchers:VEGA

- Wound Solid Propulsion motor cases; - CFRP skirts; (Avio)

CFRP Fairing (RUAG)

CFRP Payload Adapter, (Airbus DS Spain)


Building Block Approach for the Development and Verification of Composites

Structures

Limit Loads

Design Loads

Test Loads

Design Allowables

(A-values; probability 99% on a confidence level of 95%)

Strength & Life Analysis

(Positive MoS vs. Failure Criteria)


ECSS-E-ST-32-10C (Structural factors of safety for spaceflight hardware)

Applicable Factors for Composite Structures

(all factors for un-manned)

• Test factors • KQ – qualification (1.25)

• KA – acceptance (1)

• Design factors• KP – project maturity

(1.05-1.2, typical)

• KM – modeling (1.2)

• KMP – margin policy (launch vehicle specific)

• KLD – local design(1.2 in discontinuities, joints, inserts)

• Factors of Safety• FOSY (yield, N/A to

composites)

• FOSU (ultimate, 1.25, re-entry, 1.25)

• Recursive approach fromsystem to subsystem level

Coefficient Satellite Launch vehicles and pressurised hardware Man‐rated systems

Coef A or Design factor

KQ x KP x KM KP x KM KP x KM

Coef B FOSY x KLD FOSY x KMP x KLD FOSY x KLD

Coef C FOSU x KLD FOSU x KMP x KLD FOSU x KLD


Analysis and Testing of Composites Structures – Damage Tolerance

ECSS-E-HB-32-20: Structural materials handbook :• The nature and extent of analysis and tests depend upon the applicable

experience on similar structures. In the absence of experience tests of components, subcomponents and elements are performed.

• It is impractical to evaluate the damage tolerance at each location by testing. Analysis is therefore needed. The analysis procedures are based on fracture mechanics, which include stress intensity, inter-laminar fracture toughness and propagation growth rates of cracks or delamination.

• Analytical calculations are used for residual strength calculations and fatigue analyses under the loading spectrum. Sufficient testing is required to validate the analysis methods. Full-scale structural testing occurs too late to be useful during the design phase, but it is used later to update the initial damage tolerance evaluations and proceed to the final verification.

• The effects of environment is accounted for in the analysis and during the materials characterisation stage.


Composite technology developments

• Ceramic Composite structures;

• Composites for inflatable structures;

• Thermo-plastics;

• Manufacturing process developments, incl manufacturing in space (additive layer parts, pultrusion of booms)

• Enhancement of properties by doping with CNT or nano Carbon fibers;

• Grid/lattice composite structures;

• Propellant tank applications, incl liners;

• Interface design, joining technologies;

We are faced with

-More demanding performance requirements

-Need for cost reductions

Need for further development of Composite Structures Technologies!

Examples of Composite Structures R&D Projects in progress or foreseen:


Ceramic Composites

Material Technology drivers:

Why Ceramics: Ultra-low CTE (alignment/stability)

But must also have:

• High specific stiffness

• Good strength characteristics

• Good thermal conductivity

• Stability over wide temperature range

Material examples: SiC, CeSiC, HB-CeSic, Si3N4

• SiC (SiC grains bonded by sintering) well characterised, extensive flight heritage.

• CeSiC (chopped carbon fibre with Si infiltration)

• HB-CeSiC (directional carbon fibres with Si infiltration)

• Si3N4 (Silicon Nitride)


Ceramic Structures: Analysis & Verification

Current verification approach:• Deterministic;

- Derive material equivalent A-values.- Finite element analysis- Conservative safety factor ≥ 2.5- Proof testing / NDI.

Future verification approach:• Initial design is deterministic.• Subsequent deterministic/probabilistic.

- Derive material equivalent A-values.- Initial sizing deterministic.- F.E. analysis – mesh density appropriate to structure- Detail design establishes failure probability

(Project/Application dependent?).- Proof testing/NDI – to provide quantifiable improvement in

failure probability.

Ceramic specific problems:• Low fracture toughness, i.e. brittle fracture.• Susceptible to surface and volume flaws.• Low damping.


Inflatable Composite Structures

Next generation of spacecraft, both scientific and commercial will require large appendages to be deployed, e.g. => Solar Arrays (10-> 100 m2), Radiators, Sun shield, Low frequency antennae (L-Band) with less stringent surface RMS requirements….

Inflatable structures can be used for deployment and as support structure.

Main Advantage: low packaging volume and low mass, cure in orbit (thermal, UV)

Thin film solar array (Airbus DS France)

Boom deployment (Airbus DS France)


Inflatable Structures for spacecraft:Current technology developments

Driving Requirements for composite materialsfor inflatable structures:

• foldable;

• long-term storage in uncured state;

• low outgassing;

• AtOx resistance;

• rigidisable (curing) in space, preferably reversible for on-ground verification.

• low energy curing;(light,UV,thermal)

• high performance fabrics.

Verification/certification challenges:

• Consistent mechanical properties after rigidisation.

• On-ground verification. Rely on advanced non-linear numerical simulations due to complexity of testing;

epoxy resincarbon fibers +45/0/-45

thermal curing(TAS France)

epoxy resinglass fibers 0/90visible light curing(Airbus DS France)

epoxy resinglass fibersUV curing

(Airbus DS France)


In-space rigidisation of Inflatable Structures:Verification

Deployed boom (3m) + curing systembefore integration

Boom in solar simulator chamber (IABG)Before door closure

Curing achieved in 16 hours (inc. transient phases)

•New resin development•Curable and hardenable resin based on GMA-co-BuA (glycidyl methacrylate and butylacrylate).(Airbus Patent WO 2010/106017 A1).•Resin designed to have both a low outgassing and a low viscosity by controlling the molecular weight and the polymer architecture.•Composite material with Carbon or Kevlar fibers

•packaging efficiency demonstrated;

•rigidisation in In-Orbit Experiment relevant conditions

Next Step = Flight Demonstration


Antenna Composite Structure Developments

Near term technology developments:- Mainly Ka and Q/V band antennae

- More demands on reflection losses, depolarisation, thermal stability

- Automation of manufacturing processes;

- Full carbon architectures (including CFRP honeycomb core and CFRP rods)

- Reduction of thermal control hardware, therefore larger thermal range.

- Integrated metal coating

STAAR EM ultra light reflector (RUAG)

Ka band 2.4 reflector manufactured with Fibre Placement technology (Airbus DS Spain)

New concept of Dual Grid Reflector (RUAG)

Full carbon Ka band Gregorian antenna (STANT) during vibration testing (HPS-GmbH)


Antenna Structures

Longer term developments:

- CFRS material for large foldable apertures

- Surface re-shaping in orbit (actuators)

- Nano-species in the composites for improved

transverse properties (conductivity improvements)

The SMART concept and SAFIRS (ITI) demonstrator developed by Technische Universität München-Lehrstuhl für Leichtbau (TUM-LLB) [D]


Verification/Certification challenges:

• Complex kinematics analysis;

• Definition and verification of deployment margins;

• Testing on-ground complicated by air and gravity;

Antenna Structures:Verification/Certification

TALDES 6 m demonstrator of Conical ring - “V” folding (ESA Patents 568 and 621)


Carbon Nano Tubes (CNT) doped material developments

Technology drivers:- Electrical and thermal conductivity;- CTE improvements;- Damping;- Mass reduction (stiffness, strength, damping, damage tolerance)

Focus is on conductivity!

Polymers/CFRPCOPE “Composite materials for Payload and Platform Elements”

Telecommunication spacecraft carbon structure (antenna, radiator, deployable structure, wave guide)

Sample materials: HTA/T300 + LTM123,K13C2U + LY556, HTA/T300 + LY556, T700 + LY556

Improved Electrical conductivity, transverse thermal conductivity and infrared emissivity.

Demonstrator: radiator application (K13C2U + LY556) :Increase of approx. 6% with respect to the heat transfer. Test verification in progress.

NAREMA “Nanotube Re-inforced Structural Materials for spacecraft Applications”General spacecraft and launcher carbon structure improvements;

Sample materials: T700 + LY556, HTA + LY556, PEEK with 30% CF by Victrex with different types of % of CNT/CNF/GNP (0.5% and 1%) or with bohemite, shortened and carboxylated CNT and different types of manufacturing processes (pultrusion, pre-preg, CNS fabric prepreg, compounding and hot-pressing )

Additional activities foreseen in 2012 aiming a process improvements!

~25µm

microscopic image of CNT-resin sample (HPS)

Metal- and ceramic matrices:Metal matrix composites: - increase of thermal conductivity- decrease of CTE; - improvement of mechanical properties;

wear/tribology improvements in mechanisms


Carbon Nanotubes doped material developments

Main possible applications:Near-Term:Focus is on conductivity driven applications!

Key points for future technology developments:

Consolidation of the material property and process aspects,special focus on thermal and electrical property in the transverse direction;-> development of methodologies for verification of properties in as manufactured flight hardware;


High conductivity CFRP Sandwich Panel developments;

Technology drivers: Need for Highly conductive composite panels (Telecom)

“Highly Conductive CFRP Sandwich Technologies for

Platforms” study initiated.

The main purpose of this study is to design, manufacture

and test a CFRP spacecraft sandwich panel structure (for example a shear web supporting dissipative equipments) using high conductive CFRP in order to reduce the mass of large equipment panels.

– Baseline materials: K13D2U/MTM46 Pre-Preg face-skins with an aluminium core.

~25µm

Concept for Demonstrator panel (including Mass/Thermal dummies):

Testing will include; Mechanical properties (stiffness, strength), Vacuum compatibility, Thermal conductivity, Thermal gradient and Thermal cycling behaviour, Modal & Mechanical vibrations/Acoustic test, Insert pull out test and CTE evaluation.

=> Interface to Alu boxes to be addressed also!

Facesheet manufacturing


Thermoplastic Composites

Application of thermoplastic resin composite structures being re-visited:

Main driver: manufacturing process advantages! => Cost!

Technology Studies performed or foreseen:

Pre-development programmes were carried out for feasibility of out-of-autoclave processing of thermoplastic structures:

400 C processing of Carbon/PEEK with electrically heated ceramic tooling, no oven required, vacuum pressure only.

-> Secondary assembly technologies using AIB (Amorphous Interlayer Bonding ) and welding

-> Carbon/PEEK strut and bracket development. (integrated end fittings)

Planned new development activity:

Automated fibre placement processing with in-situ consolidation:

-> Fully automated one step manufacturing process.

-> Material characterization programme, demonstrator manufacturing and testing

Near term applications:

CFRP Struts for spacecraft and launchers


Composite Electronic Box Housings

Technology drivers:– Mass reduction

– Consistent CTE between box and composite support structures

State of the art:– Prototype of composite box has been designed

and manufactured using Resin Transfer Molding (RTM) method;

– Mechanical performance requirements have been verified by exposing the box to relevant vibration environments.

Key points for future technology developments:– Improvement of electromagnetic shielding, radiation shielding and thermal control

performances.

– Cost effective manufacturing of composite electronic box for small series e.g. by having available a modular metallic mould that can be easily adapted to various box sizes.

– Materials: K1100, K13C2U, K13D2U; hot cured epoxy Araldite LY556 /Aradur906/ Accelerator DY070

Follow-up studies, focus on: shielding and thermal interface


Composite Flow Moulding Components – Fasteners and Inserts

R&D Project to evaluate performance

Matrix material: PEEK

- GFRP and CFRP fasteners

- GFRP and CFRP inserts

Tests performed:- Tensile, shear and torsion strength at

different temperature

- Thermal and mechanical cycling

- Creep (72 hours)

- Relaxation (72 hours)

Follow-on project planned:

- Manufacturing and environmental testing of a demonstrator structure with CFRP fasteners

- Long term creep and relaxation behaviour

Tensile strength, M4 CFRP bolt

Icotec ag


Composite technology developments for future launcher applications

Pressurised structures

Cryogenic Propellant Tank with Thermoset Resin

Cryogenic Propellant Tank with Thermoplastic Resin

CFRP SRM

Unpressurised Structures

Advanced Interfaces

Interstage Structures

Upper stage Novel CFRP Demonstrator, Payload Adapter

CFRP Engine Thrust Frame

Fairing

Composite structure developments for future launchers are mainly performed withinthe ESA Future Launchers Preparatory Program (FLPP), Materials & Structuresprogram, supported by technology developments within other ESA R&Dprograms (TRP, GSTP);



Composite Interstages and Tanks with Thermoset Resin, (Airbus DS Spain)

Technologies developed:a) Cryogenic propellant compatibility of thermoset resin

systems (material selected M21E/IMA)b) One shot CFRP tank manufacturing (Fiber Placement)c) Permeability tests of the full tank in cryo facilityd) Non-metallic liners

Development and demonstration of composite structures technologies for launcher upper stages.

Challenges: withstand launch loads (part of the Primary Launcher structure) as well

as the pressure loads in a cryogenic environment; comply with critical leak rates, no micro-cracking; ensure thermal insulation to the surrounding system; limit the heat flux between the two compartments containing LH2 and

LOX. (prevent excessive boil-off of LH2 or sub-cooling (and perhaps even freezing) of Oxygen)

Main reason for use of composites is mass- and cost reduction!Airbus DS Spain CFRP Tank

First demonstrator during tests

Airbus DS Spain CFRP TankSecond demonstrator before tests


Cryogenic Propellant Tank with Thermoplastic Resin, (EireComposites)

Challenges:

Micro cracking leak rates of the tank.

Thermoplastic composites provide superior toughness compared to thermoset materials. This helps mitigate risk for micro cracking under cryogenic temperatures with repeated thermal cycling.

Developed Technologies:Cryogenic Upper Stage with integrated LH2 and LOX tanks:

Linerless tank is baseline design (Use of film liner after assembly backup);

Automated Tape Placement of main components (tank cylinder, domes);

Assembly using thermal fusion (induction welding) of thermoplastic composite components. As a backup solution, Amorphous Interlayer Bonding (AIB) with PEI (Polyetherimide) be considered.

Cryogenic propellant compatibility of thermoplastic resin:

Materials screening and characterization including permeability tests, IM7/PEEK (Suprem);

Impact tests in LOX environment.

Verification: Subscale demonstrator tank to be subjected to thermal environment, mechanical, pressure

and leak tests;

Permeability testing will be performed while subjected to external mechanical loading.

Induction welding

Tank Cylinder manufacture with Tape Placement

SRM: Thermoset Fiber Placement

Automation of manufacturing

From hand lay-up To automatic processesManufacturing Facility at DLR-ZLP München

Challenges



Advanced Interfaces, (SONACA)

Technologies to be matured:

Demonstration of potentiality of lightweight CFRP segmented ring made of RTM elements. By combining braids of CFRP as RTM Pre-form, an advanced, curved Y/section can be realized in one shot manufacturing.IM7 baseline fibre (IM6 or T300J as backup)RTM 6 baseline resin (Cycom 977-20 or BMI 5240-4 as backup)

Static tests on prototype

Interstage Structures (EADS-CASA)

Technologies matured:1. Co-curing of the “Omega” stringers and the skin in one

shot, manufactured by means of inflatable bags and are co-cured to the skin

2. I/F rings obtained by integrated 90 degrees fibers where the chapels are machined for the bolts installation.

Omega stringers

Lower ring: Integrated CFRP Interface

CASA ESPACIO ISSwith co-cured omega stringers


Upper Stage Novel CFRP Demonstrator (EADS-CASA)

Based on existing Adaptor technologies available at EADS-CASA, further development activities are performed to optimize cost efficiency. A strong focus is on automated processes for manufacturing to reduce manufacturing effort and secure quality. Material used is M40J/977-2.

Technologies to be matured:

a. Single shear upper interface (replacing the previous double shear)

b. CFRP joint (I/F rings): I/F obtained by an integrated CFRP ring produced automatically in the FP machine by laminating circumferential (90º) layers (replacing the previous RTM rings reducing costs for inspections and manual work).

c. Optimization of design for simplified, fully automated manufacturing.

The development includes a full scale, 2624 mm Launch Vehicle interface, demonstration in manufacturing and qualification-like static test program.

Upper I/Fsingle shear config.

Lower I/Fintegrated CFRP rings

Conical adapter


Grid/Lattice –structures:Verification/Certification

Developments addressing:

- Manufacturing technologies (Winding and Fiber-Placement);

- Node design (as different technics are available, from simple crossed node to so-called steering nodes, or alternatively cut nodes);

- Interface design (metallic or CFRP ring)

- Overflux aspects;

- Shock load attenuation;

Verification/certification challenges mainly related to control of manufacturing quality.

Focus on launcher adapters and stable structures for spacecraft;

Targeted benefits: mass, cost, shock load attenuation!

Cylindrical demonstrator fully representative of the lower part of the

final VEGA adapter(in terms of grid pattern and rib length)

tested during a TRP activity

Cylindrical demonstrator tested at Airbus DS Spain for validating the node design:• Rupture predicted at around 600kN (which is already twice what we

would need for the complete VEGA adapter ), prediction based on flat sample tests and associated correlation.

• During the tests 945kN (96t) was achieved.

96t were withstood by a lattice structure weighting less than 10kg!!!


CFRP Engine Thrust Frame, (Dutch Space)

Technologies to be matured:– CFRP Technology in Thrust Frame Application.

– Equipment arrangement and fastening, Interface design and Analysis.

– Conical shell will be manufactured using Automated Fiber Placement technique, M46J.

– End Cap, this is the interface structure between the cone and the engine, will be manufactured using RTM with integrated CFRP actuator lugs for TVC actuators.

Challenges

– High thermal gradients (90 - 350 K) in combination with CTE incompatibility to metallic cryo tank.

– Highly loaded interfaces; Engine loads: 400 kN, Actuator loads: 100 kN

Static Test Campaigns for demonstrator;

Cone Manufactured with Fiber SteeringThermoplastic Conical Stiffened Skin panels with stiffeners integrated in the lay-up and

co-consolidated.

End cup manufactured with RTM


Fairing, RUAG (CH)

Fairing concepts and manufacturing processes are developed with the aim to be more cost effective and improve performance.

1. New and/or alternative manufacturing methods (Low pressure resin infusion moulding) are evaluated with the focus on efficiency and reduction of internal interfaces/junctions.

2. Alternative materials are investigated and subjected to trade-off (Monolithic or Skin/Stringer) Honeycomb core with facesheets of Carbon Fiber UniDirectional/175°C Curing epoxy).

3. Increase temperature limit of sandwich facesheet could enable a reduction of TPS mass on the fairing, contributing to overall lower mass and simplified recurring production;Possible use of BMI to be studied;

4. Acoustic performance improvements:– Increase circumferential stiffness/mass ratio with

stiffer fibre and new lay-up, to increase ring frequency from 100 to > 300 Hz

– Helium filled fairing

Half of the VEGA Fairing• OoA• elimination of joints


Increase of Bolted Joint Performance

Objectives:a. Performance improvement of highly loaded bolted composite structural jointsb. Demonstration by numerical models, sample testing and component testing;c. Demonstration of selected concept on breadboard models;d. Contractors:DLR, HPS, EADS CASA Espacio, Inegi;

Users: MT Aerospace, Kayser-Threde,RUAG, SABCA.

Technology: CFRP locally reinforced with titanium metal foils

Results:a. Significant improvement in

bolted joint performance for most applications Significant mass savings; (several hundred kilo mass saving for an Ariane 5 solid booster segment)

b. Accurate numerical predictions are possible;

c. Feasibility of manufacturing and inspection in an industrial context shown. Process improvements are needed.

Demo adapter, FP with Ti foils integrated (EADS-CASA)



Increase of Bolted Joint Performance for CFRP

Material Dependability/Availability:

CFRP materials are used in every space structure: satellite and launchers.

Spacecraft users are pushing the stiffness and strength performance frontier, but not the quantity; (about 0.1% of total consumption in Europe)

CFRP suppliers consolidation has been seen in the last years, where the remaining “European” suppliers seems to be driven by “volume market” industries: Aeronautics, automotive etc.

The PAN UHM fibers are a monopoly from a Japanese industry.

Some semi-products (fabrics / pre-pregs) are subject to ITAR restrictions;

Problem for space applications?

Availability;

Long lead time;

Risk of obsolescence;

Solution? 1) European suppliers,

2) Use of more readily available material systems “aeronautic”

However: Change of material => costly delta qualification for each structure;


Material Dependability: solution example for semi-product!

“European CFRP Honeycomb Material”,Contractors: INVENT (D), HPS (D)

Objectives:- Development of a competitive CFRP honeycomb

(price, mechanical performance, delivery times) with European materials;

- Qualification for application in stable structures like optical benches, reflectors, etc.

Status:- Materials SUCCESFULLY manufactured with European and Japanese fibers)- Material testing completed;- Demonstrator (reflector) manufactured; Two antenna projects have already used the Carbon Core of Invent: STANT multiple optics earth-deck antenna (Ka band, multibeam application): successfully tested for

RF and environmental loads. Ka band Dual Gridded Reflector supporting cylinder: successfully tested for environmental and RF.


Composite Structures:Analysis tools and methodologies

Identified need for improvements in analysis tool and methodologies

Damage tolerance verification for new developments;

• e.g. Assessment of delaminations;

Evaluation of production damages (e.g. impact damage)

Focus on industrial application


Examples: low velocity impact damage

#falling objects # Visible external dent of ~Ø25 mm.

Internal damage (core

crushing, core/facing

debonding, and

delaminations) affecting

Ø80 mm

EADS CASA


Delamination Assessment Tool for Composites Development

• Development of tools and methodologies for damage tolerance assessment of composite structures.

• Complementing available standards and best practice manuals with regards to damage tolerant design

• Damage tolerance verification methodology including analytical and numerical methodologies, material and component testing, NDI

(HPS Lda. PT)

Damage Tolerance Methodology


Analysis Methodologies:Delamination tool development

1. Development of a methodology to address delaminations in composite structures.Prime contractor: HPS (Portugal)

2. Guidelines are prepared for different analysis methods, and applied in a demonstrator:

a. Level 1: prediction of delamination stability by application of a Crack Tip Element (CTE)

b. Level 2: prediction of delamination stability by application of the Virtual Crack Closure Technique (VCCT)

c. Level 3: prediction of delamination stability by application of Cohesive Finite Elements.

3. The main activity focuses on delaminations, but exploratory work on impact damage is performed as well, which will be continued in a follow-on activity.

4. The main activity will be finished mid-2015.Future Trends in Certification of Advanced Technology Structures, Bristol 16 September 2015

Design of Composite Structures with ESAComp Software

ESAComp Facts1. Developed under ESA Contract, maintained and

distributed by Componeering (http://www.componeering.com)

2. Original development by Helsinki University of Technology also under ESA contract

3. Current release: ESAComp 4.5.24. Yearly ESA maintenance contracts

ESAComp Key Features1. Graphical user interface. 2. ESAComp objects, multi-level database and Data Bank. 3. Multiple analyses and graphic display of results. 4. FE export/import interfaces. 5. User extensions. 6. Comprehensive documentation.

ESAComp Analysis Capabilities1. Fiber/matrix micromechanics 2. Analyses for constitutive and thermal/moisture expansion behavior

of plies 3. Laminates

a. 2.5D behavior - classical lamination theory (CLT)b. parameterized "theta-laminates" and "p-laminates" c. load response failure and design envelopes d. probabilistic 2.5D behavior e. sensitivity studies for 2.5D behavior and FPFf. notched laminate analysis of circular and elliptic holes

4. Panels a. flat and curved panels with or without stiffenersb. Mindlin plate analysis using integrated Elmer FE solverc. rectangular plates d. load response, failure and stability under applied loads e. natural frequencies

5. Beams and columns 6. Bonded joints 7. Mechanical joints

a. single and double lap joints under axial loadsb. fastener and by-pass loads, laminate stresses and

strains on fastener holes, margins of safety for failure, prediction of failure mode



ESA Space Science ProgrammeROSETTA and PHILAE Success

PHILAE’S INSTRUMENTS

Philae’s Drill

Support Truss

Base Plate

Instruments Carrier

Solar Hood PHILAE’s STRUCTURAL CONFIGURATIONA structure consisting of lightweight CFRP sandwich plates and CFRP + Kevlar frames and rods was chosen in order to achieve the required high stiffness and strength with the given mass budget. The overall mass of the primary structure is 18.2 kg.

REDUNDANT PHILAE’s ANCHORING SYSTEMIt consists of two pyrotechnically actuatedAnchoring Harpoons and a redundant Control Electronics

Rosetta is an interplanetary spacecraft whose main objective is to rendezvous with Comet 67P/Churyumov-Gerasimenko. In order to investigate the comet nucleus and the gas and dust ejected from the nucleus as the comet approaches the Sun, Rosetta carries a suite of eleven instruments on the comet orbiter and Philae, a lander equipped with further ten instruments which perform surface measurements.The orbiter instruments combine remote sensing techniques, such as cameras and radio science measurements, with direct sensing systems such as dust and particle analyzers.

Rosetta's lander Philae has woken up after seven months in hibernation on the surface of the Comet!!


Some additional remarksKnowledge

Major elements in work environment: People/Processes/Technology

People: Know how & experience, communication & perception,collaboration, cultures, global environment, etc.

Processes: Sharing knowledge, reusing knowledge, best practices, etc.

Technology: Knowledge in maths & physics, engineering, computing hardware & software, data management, etc.

Knowledge plays a key role throughout. However, knowledge is not a

privilege; it has to be earned through hard work.

Mrs. Tiziana Cardone email: [email protected]

Mr. Julian Santiago Prowald email: [email protected]

Mr. Torben Henriksen email: [email protected]

Mr. Rafael Bureo Dacal email: [email protected]

Prof. Dr. Constantinos Stavrinidis email: [email protected]

Possible contacts for this presentation:


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