DGLR / CEAS European Air and Space Conference 2007
Space Transportation Systems Propulsion and Structures
Status of Discussion of the DGLR Expert Committee DGLR-Fachausschuss S4.1
R. Lo (AI), W. Zinner (Astrium RT), R. Pernpeintner (MT Aerospace ) DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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DGLR / CEAS European Air and Space Conference 2007
Introduction •
The DGLR S4.1 Working Group on Space Transportation Systems (Fachausschuss S4.1 Raumtransportsysteme) is a forum for members of agencies, institutions, industry and universities. Gathering and analysis of information, argumentations about past, present and future space transportation systems are the objectives of this particular group.
•
Analysis and documentation is coordinated around the topics – Demand & Market – System Concepts & Subsets – Propulsion, Structures & Subsystems (System related aspects)
– Missions & Operations (incl. ground infrastructure) – Cost (Development, Production & Operation) – Projects / Programmatic (Development & Demonstration) •
This paper presents the status of information and analysis of DGLR-FAS S4.1 about propulsion and structures.
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Content of Presentation Jet and Rocket Propellant Classification Propellants in Use / Propellants of Current Interest Environmental Benignity: TEHF+Green Hybrids Design Examples of Liquid Rocket Engines Structures of Liquid Propulsion Stages Design Examples of Solid Rocket Motors Structures of Solid Propulsion Stages Design Examples of Hybrid Rocket Motors Structures of Hybrid Propulsion Stages Propulsion, Structures and Subsystems - Results and Facts Summary and Outlook
Σ: Propellants + Liquid-, Solid-, Hybrid-Propellant Engines
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Jet Propellant Classification and Consumption
Propulsion type
Characterisation
Remarks
Airbreathing:
Mixed hydrocarbons
Rocket:
LOX/LH2
< fully cryogenic
3-4000 t
LOX/HC
< semi cryogenic
9500 t
N2O4/Hydrazines
< hypergolic storable
10000 t
N2O/Polymere
< green storable
10-50 t**)
AP/HTPB/Al
< Solid storable
9000 t
Jet propellant / Aviation fuel
~ Annual consumption*) 53,91 Mio.t / 0,897 Mio. t
*) Rockets: estimated average 2006/6-2007; **) SS1 2003-2004 SL tests + 6 flights
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(AI) Overview of Propellants in Use Propellants currently used in operational space systems: Number of stages with individual propellants launched 1/2006 to 6/2007: LOX/ LH2
LOX/ Kero
6
0
30
0
2
14
0
0
0
14
India
1
0
2
4
20
Japan
10
0
0
0
18
USA
13
20
11
0
74
Russia
0
104
85
0
5
Ukraine
0
18
15
0
5
China
EU
DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
NTO/ UDMH
NTO/ UH25 MMH
Solids
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(AI) Propellants of Current Interest Some physical and chemical liquid propellant characteristics: Prop.
Fp °C
Bp °C
Sol.Ds.
Liqu.Ds.
Rem. + (Isp 68:1)
LH2
-259,2
-252,8
0,088
0,0711
+24% „100% slush“ (391)LOX
CH4
-182.5
-161.5
0.466
0.423
Tcrit = -82,7°C (311)LOX
C3H8
-190,0
-42,1
0,582
7,1 bar VP at 20°C (~305)LOX
O2
-218.8
-183.0
1,14
TEHF= MpMission/L50 = 0
N2O4
-9.3
21.15
1.45
TEHF = 11-14
H2O2
-0.4
150.2
1.44
TEHF = 3-5
RP-1
~-40
177-274
0.820
TEHF = 0,09; (300) LOX
N2H4
1.4
113.5
1.004
TEHF = 2,8; (292)N2O4
63.0
0.793
TEHF=3; (287)N2O4
UDMH -57.0
1.46
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(AI) Environmental Concerns: Green Propellants/greenHybrids Prop.
Fp °C
Bp °C
N2O
-90.8
-88.5
Polyethylene
~110
(477)
Density Fp kg/l
0,92-0,96
Density Bp kg/l
Rem.
1.22
Tcrit: 36,6°C Pcrit: 7,27MP TEHF = 0 (254)HTPB
n.a.
20°C
• Environmental concerns refer to storage and handling rather than emissions ! • Comparison of space transportation emissions (100 launches/a) with air traffic pollution and terrestrial traffic yields 1 : 2000 : 5500 • HCl emission of solid rockets : coal combustion = ~1/200 • In abs. numbers: solid rockets = 0,01MT HCl/a; coal combustion:1,8MT; volcanoes: 7,8 MT; oceans: 300MT
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Design Examples of Liquid Rocket Engines Basic Principle of a Liquid Propulsion System
Fuel/Ox Tank
Turbopump Combustion Chamber Nozzle
Pressurization System: ¬ Pressure- or turbopump-fed systems raise the pressure above the operating pressure of the engine Thrust Chamber Assembly: ¬ Generates thrust by efficiently converting the propellant chemical energy into hot gas kinetic energy
Thrust
The Thrust Chamber is the Heart of All Liquid Propellant Rocket Engines DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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Design Examples for Liquid Rocket Engines Non-Storable (cryogenic) Propellants
Liquefied gases stored at very low temperatures Liquefied to minimize the sizes of tanks needed to store them Most common propellant used for today's rocket applications are: ¬ Fuel: Liquid hydrogen (LH2) ¬ Oxidizer: Liquid oxygen (Lox)
Pros ■ Lox/H2 provides the highest specific impulse (~450 sec) ■ Lox/H2 is environmentally friendly ■ Non corrosive Cons Thermally insulated tanks Prior to loading, tank evaporation to avoid frozen particles Venting on the launch pad and refilling Low temperature design Expensive DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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DGLR / CEAS European Air and Space Conference 2007
Design Examples for Liquid Rocket Engines Storable Propellants
Distinction between earth and space storable propellants Liquid at environmental conditions (earth or space) Storable for a long time in sealed tanks Most common propellants for today's rocket applications are: ¬ Fuel: Kerosene, MMH, UDMH, N2H4 ¬ Oxidizer: N2O4
Pros ■ Stable at ambient temperature and pressure, i.e. no boil-off ■ Non reactive with tank materials ■ Instant readiness of the rocket engine Cons Medium performance Hypergolics are extremely toxic (N2O4/MMH) Surface contamination High handling safety precautions DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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DGLR / CEAS European Air and Space Conference 2007
Design Examples for Liquid Rocket Engines Engine Cycle
The engine cycle (thermodynamic cycle) terminology refers to the source of energy to drive the engine's turbines.
Cycle Analysis Selects the proper thermodynamic cycle, e.g. open versus closed Delivers: ¬ Engine operating parameters (Isp, thrust, mass, etc.) ¬ Engine component parameters (temperatures, pressures, mass flow, flow areas, etc.) ¬ Concept trade-offs Types of Engine or Power Cycles Up to now, only four cycles have been developed and flown: ¬ Pressure-fed -, gasgenerator -, expander -, staged combustion cycle ¬ The gasgenerator cycle was the first in use One further cycle is currently part of a US R&T demonstration program (IPD) ¬ Full flow staged combustion cycle (FFSCC) DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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Design Examples for Liquid Rocket Engines Cycles of Flight Engines Pressure-Fed Cycle
No pumps/turbines Low thrust Pressurized tanks feed the propellants into the main chamber Simpler engine design
Gagenerator (GG) Cycle
Open cycle Medium to high thrust A small amount of the propellant is burned separately in the GG to drive the turbines Turbine drive gas is pumped over board and not routed back to the main injector
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DGLR / CEAS European Air and Space Conference 2007
Design Examples for Liquid Rocket Engines Cycles of Flight Engines Expander or Bleed Cycle
Closed or open (bleed) cycle High performance/medium thrust No gasgenerator/preburner All or a portion (bleed) of cooling channel gas is used to drive turbines
Staged Combustion Cycle
Closed cycle Medium to high performance/thrust Propellant is burned in two stages: preburner and main chamber All propellant is mixed and burned in the main combustion chamber (MCC)
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Design Examples for Liquid Rocket Engines Cycles of Flight Engines Full Flow Staged Combustion Cycle
Closed cycle High performance/medium thrust All propellants pass the turbines Lower turbine gas temperature by using a full flow cycle Longer engine life Challenge: Oxidizer-rich preburner Most efficient rocket engine cycle Synthesis/Conclusion
Europe combined the high energetic propellant (Lox/H2) with the medium efficient gasgenerator cycle US (SSME), Russia (RD-0120) and Japan (LE-7A) have already applied the high energetic Lox/H2 to the high efficient staged combustion cycle The technology step combining Lox/H2 with staged combustion might be the next generation engine in Europe DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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Solid Rocket Propellant- and Motor-Classification: conventional
Motor/grain shape: > Cylinder grain with Spherical Solid internal star cross Rocket Motor (Historic Conventional Black Arrow Waxwing) section (below) propellants: v Storable „conventional“ AP/HTPB/Al grain (right): Endburner: Spherical kickVery high stage motor, incharge density space applications (good for HEDM!) (above) DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems 15
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Solid Rocket Propellant- and Motor-Classification: cryogenic
Grain design: > Cryogenic solid propellants CSP V
Modular cylinder grain:
Quasi homogeneous alternating stack of ox.- cylindrical sponge and fu-moduls; igniter, grain gasgenerator (below)
CSP use frozen liquids, e.g. based on solid hydr. peroxide or oxygen with polymers. Modular end-burners with Rod-in-matrix, tube bundle and concentric layer design (right) DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
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(MT) Design Examples for Solid Rocket Motors
In-Orbit Motors Used for: position / apogee motors Size: Ø 4 inch – 1300 mm monolithic casing Material: steel and Titanium, also C- and Aramide-fibres Pressure: up to 6 MPa
Small In-line / Strap-on Booster size: Ø up to 1,6 m monolithic steel (roll-and-weld) or CFRP cases (filament wound) Performance factors steel: 2,7- 8,5 km CFRP: up to 14 km
Solid propellant stages size: up to 350 to propellant Ø up to 3,5 m segmented cases high strength steel or CFRP thrust vector controlled
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(MT) Design Examples for Solid Propulsion Stage Structures Common features
Propellant
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(MT) Design Examples for Solid Propulsion Stage Structures steel segments welded from individual cylinders and domes
Composite casings
VEGA P80 at AVIO
Ariane 5 MPS segment S1: Infiltration technique
Clevis-Tang Intersegment Joint
¾ retaining „nose“ against opening under pressure ¾ O-ring sealing
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DGLR / CEAS European Air and Space Conference 2007
Load Introduction Structure – Ariane 5 Front Skirt Diameter Total height Mass Load capacity Material Thermal protection
5 400 mm 3 288 mm 1 785 kg 6 000 kN Aluminum alloy, CFRP PROSIAL®
The EPC Front Skirt (JAVE-C) is located at the top of the cryogenic main stage (EPC) of the Ariane 5 launcher and transmits the thrust of the solid boosters into the central body of the launcher and equalizes it regarding payload compatibility
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DGLR / CEAS European Air and Space Conference 2007
Large Liquid Propellant Tank Structures – Ariane EPC Domes •
Comparison with gore panel method advantages of concave spin-forming: – less parts to manufacture Æ reduced costs, logistics easier – T8 of complete dome plate welds exhibit T8 condition, too Æ weight saving
Dome with gore panel method Spin-formed dome
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DGLR / CEAS European Air and Space Conference 2007
Large Liquid Propellant Tank Structures – Domes for H II-A
MT Aerospace AG manufactures for Mitsubishi Heavy Industries the 1st stage and LRB tank bulkheads of Japans launch vehicle H-IIA. The elliptical bulkheads are spinformed of one Al 2219 plate to final T8 temper. MT's spinforming technique leads to tank bulkheads with less weight, better material properties and lower cost than the conventional method used for H-II. Height
738 mm (29.1 in)
Material
Al 2219 T8
Propellants
LOX/LH2
Mass
230 kg (507.1 lb) (LOX) 180 kg (396.8 lb) (LH2)
Pressure
4.3 bar (62.4 PSI) (LOX) 3.4 bar (49.1 PSI)
DGLR Fachausschuss Raumtransportsysteme S4.1 / Propulsion, Structures and Subsystems
Wie semann 04
3 860 mm (12.7 ft)
NASDA
Diameter
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CMC Elements of Re-Entry Vehicles - X38 and CRV Heritage
nose cap
wing leading edge
high temp. bearing
„chin panel“
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DGLR / CEAS European Air and Space Conference 2007
CMC Elements for Re-Entry Vehicles - X38 and CRV Heritage
XPERT hot metal TPS configuration with ODS metal cone, CMC nose and flaps (Image Courtesy Dutch Space - EADS
X38 Hypervelocity re-enty. (Img.:NASA)
Astrium)
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DGLR / CEAS European Air and Space Conference 2007
Design Examples of Hybrid Rocket Motors
•
SpaceDev SS1 Hybrid Motor Development 2003 / 2004 (Poway, Cal.USA)
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DGLR / CEAS European Air and Space Conference 2007
(AI / ASTRIUM) Design Examples of Hybrid Rocket Motors
(above) SpaceDev Hybrid Propulsion Module Test Facility / (right) “Dream Chaser” SD Hybrid Suborbital Plane (Img. Courtesy SpaceDev)
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Propulsion, Structures and Subsystems - Results and facts • • • • • • • •
Chemical rocket propellants will be used for the next 50 years + !! Specific energy of available propellants will increase tenfold within the next 20 years (Isp triples) Environmental benignity will be the most important design driver of new propulsion systems within 5 years Liquid Rocket Engines will be used where-ever throttling and reusibility are prime requests . Solid Rocket Motors will go cryogenic within 15 - 20 years (HEDM) Hybrid Rocket Motors will continue to be used wherever their chaotic combustion is not a concern Structures of liquid propulsion stages will see very high degrees of integration Solid Propellant stage structures will need adaptation to reusibility.
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Propulsion, Structures and Subsystems - Results and facts Subsystems still to be investigated: • Integrated solid propulsion for winged vehicles • Replaceable units for propulsion, in particular ORPUs • HEDM additive production and application • Soft boundary layer (low noise) HEDM rocket nozzles • Lunar infrastructure (+ transportation): design to specific structures • ....
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Summary and Outlook
Summary and Outlook (1) • Annual launch numbers of stages with LOX-LH2 / Solids / LOX-Kero / NTO-Hydr. = 1 : 3,14 : 3,22 : 3,36 • Environmental loads caused by space traffic are negligible compared with aircraft emissions, but are unique above ~12km • Propellant handling- and environmental risks can be evaluated by a characteristic hazard number (TEHF) • Liquid propulsion is a mature technology but still able to see some further improvements towards advanced combustion cycles • Conventional storable quasi-homogeneous solid propellant grains will be replaced by cryogenic modular ones in all missions without long term storability requirement
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Summary and Outlook
Summary and Outlook (2) • • • •
Ceramic Matrix Composites are a key technology for hot re-entry structures Advanced production methods can reduce costs (e.g.: spin formed tank domes) Hybrid propulsion is fashionable but will never make it without active regression rate control One last time: high thrust propulsion will remain chemical for all foreseeable future Thats all folks! Thank you for your attention!
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