SDI Orbital Launch Systems Catalog [DO NOT POST]

[The Technocratic Syndicalists]

Hades V

General Characteristics:

• Function: Super-Heavy Lift Cargo Launch Vehicle
• Height: 116 m (with payload fairing)
• Diameter: 12 m
• Mass: 3,700,000 kg
• Payload to LEO: 200,000 kg
• Stages: 2

1st stage - Hades V Reusable Booster

• Length: 51.9 m
• Diameter: 10 m
• Empty mass 130,000 kg
• Gross mass: 2,930,000 kg
• Engines: 9x Syndicate Dynamics LR-98
• Thrust: 42,570 kN (sl), 45,270 kN (vac)
• Specific Impulse: 310 sec (sl), 340 sec (vac)
• Fuel: LOX/RP-1

2nd stage - Hades V Cryogenic Upper Stage

• Length: 26.7 m
• Diameter: 10 m
• Empty mass 54,000 kg
• Gross mass: 680,000 kg
• Engines: 2x Syndicate Dynamics LR-107
• Thrust: 6,600 kN (vac)
• Specific Impulse: 460 sec (vac)
• Fuel: LOX/LH2

First Stage:

• Name: Syndicate Dynamics LR-98
• Type: Liquid-fuel engine
• Propellant: LOX/RP-1
• Mixture Ratio: 2.7:1
• Cycle: Oxidizer-rich staged combustion
• Length: 3,800 mm
• Diameter: 2,200 mm
• Dry Weight: 2,580 kg
• Nozzle Ratio: 40:1
• Thrust (vac):5,030 kN
• Thrust (sl): 4,730 kN
• Throttle Range: 100% to 65%
• Thrust-to-weight ratio: 190
• Chamber Pressure: 19.3 MPa
• Isp (vac): 335 seconds
• Isp (sl): 300 seconds

Second Stage:

• Name: Syndicate Dynamics LR-107
• Type: Liquid-fuel engine
• Propellant: LOX/LH-2
• Mixture Ratio: 6.0:1
• Cycle: full-flow staged combustion
• Length: 4,500 mm
• Diameter: 2,400 mm
• Dry Weight: 4,050 kg
• Nozzle Ratio: 86:1
• Thrust (vac):3,300 kN
• Thrust (sl): 2,950 kN
• Throttle Range: 100% to 65%
• Thrust-to-weight ratio: 83
• Chamber Pressure: 20 MPa
• Isp (vac): 450 seconds
• Isp (sl): 395 seconds

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reserved

[The Technocratic Syndicalists]

Hyperion SSTO

General Characteristics:

• Function: Reusable single-stage-to-orbit launch vehicle
• Height: 48.5 m
• Width: 54.0 m
• Mass: 1,206,500 kg
• Payload to LEO: 45,000 kg
• Payload to GTO: 20,500 kg
• Stages: 1 or 2

1st stage - Hyperion SSTO

• Empty mass 96,500 kg
• Gross mass: 1,161,500 kg
• Engines: 7x SDI LR-3300
• Thrust: 15,400 kN (sl), 18,200 kN (vac)
• Specific Impulse: 385 sec (SL), 470 sec (vac)
• Fuel: LOX/LH2

2nd stage (optional)- SDI Algol-H

• Empty mass 3,000 kg
• Gross mass: 19,500 kg
• Engines: 3x SDI LR-200
• Thrust: 210 kN (vac)
• Specific Impulse: 490 sec (vac)
• Fuel: LOX/LH2

Design & Construction:

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The Hyperion is a VTHL (vertical take-off and horizontal landing) SSTO (single-stage-to-orbit) reusable launch system designed by SDI Aerospace Systems.

Design & Construction:

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The Hyperion employs a high L/D ratio lifting body design which is designed to reduce thermal loads and surface temperatures during atmospheric re-entry. The fuselage of the Hyperion employs a cold, integral structure design with the composite cryogenic fuel tanks acting as integral load bearing members members. The internal structure of the vehicle including the cryogenic propellant tanks is constructed almost entirely from graphite/epoxy composite with the exception of the engine thrust structure which is constructed primary from 2095 HT72 aluminum–lithium alloy. The vehicle has three propellant tanks including a forward dual lobe composite LOX tank and twin quad lobe LH2 tanks located on either side of the payload bay which are constructed from graphite/epoxy composite with stainless steel anti-slosh baffles and an outer layer of polyimide foam (APF) insulation. Propellant feed lines are constructed from graphite/epoxy composite with stainless steel bellows and 6Al-4V titanium alloy restraints and an outer layer of foam insulation. Unlike the fuselage the wings of the vehicle feature a hot structure design with a titanium metal matrix composite inner structure consists of ribs and spars constructed from a silicon carbide (SiC) fiber reinforced rapid solidification rate (RSR) 6Al-4V titanium alloy matrix.

The windward surface of the vehicle (front, sides and bottom of the fuselage) along with the leading edges of the wings employ a ceramic composite thermal protection system (TPS) consisting of an outer layer of carbon fiber reinforced silicon carbide (C/SiC) panels attached to the underlying vehicle structure with C/SiC standoff posts with the air gap between the polyimide foam (APF) insulation and the C/SiC panels being filled by multiple layers of internal multiscreen insulation (IMI), a sandwich of platinum and gold sheets separated by ceramic spacers. The lightweight ceramic thermal protection is intended to withstand leading edge temperatures exceeding 1600°C while being reusable and not needing maintenance or replacement between missions. The control surfaces of the vehicle including ailerons, rudders, and body flaps feature an entirely C/SiC construction with C/SiC skin panels and internal structure fastened together with Inconel 718 alloy bolts. The leeward surfaces of the vehicle, which do not experience nearly the thermal loads of the windward surface and leading edges, features a tailorable advanced blanket insulation (TABI) thermal protection system consisting of aluminoborosilicate fiber fabric based insulation which directly bonded to the APF insulation of the cryogenic tanks with silicone rubber adhesive (RTV).

Propulsion:

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• Name: SDI LR-3300
• Type: Liquid-fuel engine
• Propellant: LOX/LH2
• Mixture Ratio: 7:1
• Cycle: Hybrid gas generator/full-flow staged combustion
• Length: 4,300 mm
• Diameter: 6,400 mm
• Dry Weight: 2,075 kg
• Nozzle Ratio: 170:1
• Thrust (vac):2,600 kN
• Thrust (sl): 2,200 kN
• Throttle Range: 100% to 50%
• Thrust-to-weight ratio: 128:1
• Chamber Pressure: 22.4 MPa
• Isp (vac): 470 seconds
• Isp (sl): 385 seconds

The Hyperion vehicle is powered by a total of seven SDI LR-3300 Liquid-fuel rocket engines which are fueled with slush LH2 fuel and subcooled LOX oxidizer. The LR-3300 is a linear aerospike engine with two rows of inner and outer combustion chambers on either side of a wedge-shaped truncated two-dimensional nozzle which acts as the expansion surface of the engine. As both sides of the engine are open to the atmosphere the nozzle is naturally altitude-compensating with the exhaust plume widening and the expansion ratio of the engine increasing with decreasing atmospheric pressure as the vehicle climbs to orbit. The effective area of the expansion ramp is also increased by pumping the secondary exhaust from the gas generators through slits in the base of the truncated nozzle, a feature which also reduces the base drag of the vehicle. Unlike traditional rocket engines the LR-3300 engines are not gimbaled and are instead rigidly affixed to the fuselage thrust structure with pitch controlled by differential throttling of the top and bottom combustion chamber rows, yaw controlled by differential throttling of engines on opposite sides of the vehicle, and roll controlled by differential throttling of both top and bottom combustion chamber rows on opposite sides of the vehicle.

The power cycle of the LR-3300 is a unique hybrid gas generator and staged combustion cycle with 13 dual-combustors on either side of the truncated two-dimensional nozzle. The inner row of combustors on each side of the expansion ramp employ a full-flow staged combustion cycle with fuel-rich and oxygen-rich preburners while the outer row of combustors employ two fuel-rich gas generators. Both staged combustion preburners and gas generators employ a total of four (per engine side) turbine driven hydraulic high pressure pumps including two LH2 and two LOX high-pressure pumps each of which also had adjacent low-pressure jet pump. Throttling of top and bottom combustors is achieved via a throttling valve in the oxidizer feed line to each row of combustors. The engine is designed with modular top and half modules each consisting of gas generator and staged combustion turbomachinery sets, twin combustion chamber rows, and one expansion ramp surface forming half of the truncated two-dimensional nozzle. The exhaust flow from the top and bottom gas generators along with small amounts of bleed gases from the hot-gas manifolds in the staged combustors are used to pressurize the base of the truncated expansion ramp, increasing the effective expansion ratio of the engine and reducing the vehicle's base drag at lower altitudes. Engine cooling is achieved by pumping all the pressurized hydrogen fuel that leaves both gas generator and staged combustion turbopumps through a series of parallel up-pass circuits which cools the four sides of each combustion chamber and a single-down pass circuit which cools both sides of the nozzle. Both the combustion chambers are constructed from 3-D reinforced C/SiC (carbon fiber reinforced silicon carbide) formed using chemical vapor infiltration (CVI) with integrally machined cooling channels for the hydrogen fuel coolant.

Second Stage (optional):

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• Name: SDI LR-200
• Type: Liquid-fuel engine
• Propellant: LOX/LH2
• Mixture Ratio: 6:1
• Cycle: Expander cycle
• Length: 3,910 mm (nozzle extended)
• Diameter: 1,980 mm (deployed)
• Dry Weight: 197 kg
• Nozzle Ratio: 1300:1 (deployed)
• Thrust (vac): 70 kN
• Throttle Range: 100% to 10%
• Thrust-to-weight ratio: 37:1
• Chamber Pressure: 13.7 MPa
• Isp (vac): 490 seconds

[The Technocratic Syndicalists]

Taranis

General Characteristics:

• Function: Two-stage-to-orbit launch vehicle
• Height: 30.5 m
• Width: 19.0 m
• Mass: 174,000 kg
• Payload to LEO: 2,000 kg
• Stages: 2

1st stage -

• Empty mass 31,700 kg
• Propellant mass:94,800 kg
• Total mass: 150,300 kg
• Engines: 1x SDI RM 2100
• Thrust: 2,050 kN (sl), 2,130 kN (vac)
• Specific Impulse: 380 sec (SL), 450 sec (vac)
• Fuel: LOX/LH2

2nd stage

• Empty mass: 4,700 kg
• Propellant mass:17,600 kg
• Total mass: 23,700 kg
• Engines: 1x SDI RM 5800
• Thrust: 85 kN (vac)
• Specific Impulse: 360 sec (vac)
• Fuel: LOX/RP-1

Design & Construction:

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Taranis is a partially reusable two stage to orbit launch vehicle system designed by SDI Aerospace Systems. The Taranis consists of a reusable VTHL (vertical take-off and horizontal landing) flyback booster coupled to an expendable upper stage and is intended to act as a low launch cost, short turnaround time launch system for placing small satellite payloads into low earth otbit. The reusable first stage is launched vertically like a conventional rocket and flies on a suborbital trajectory, accelerating the system to a first stage burnout velocity of around 3 km/s (Mach 10) where following second stage separation the booster reenters the atmosphere and glides back to the launch site where it then lands horizontally like an conventional aircraft.

Design & Construction:

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The Taranis launch system consists of two stages, a reusable flyback booster and an expendable upper kick stage for placing payloads into orbit. The reusable first stage is launched vertically like a conventional rocket and flies on a suborbital trajectory, accelerating the system to a first stage burnout velocity of 3 km/s (Mach 10) at an altitude of around 60 kilometers where the second stage separates and accelerates the payload into orbit. The booster then coasts along a ballistic trajectory, reaching an altitude of around 120 kilometers where it then descends and reenters the atmosphere before gliding back to the launch site where it then lands horizontally like an conventional aircraft. The first stage features a composite warm-structure construction with a composite load bearing structure and integral stand-off ceramic matrix composite based thermal protection system which protects the vehicle structure from extreme temperatures experienced during hypersonic reentry into the earth's atmosphere. Most of the first stage including the thrust structure, wings, and propellant tanks is constructed from graphite/polyamide composite costing IM7 carbon fibers woven into a high temperature AFR-PE-4 polyamide matrix material. The nose and leading edges of the first stage are covered by a standoff ceramic matrix composite thermal protection system consisting of carbon fiber reinforced silicon carbide (C/SiC) panels which are attached to the underlying vehicle structure using C/SiC standoff posts, the air gap between the vehicle structure and the C/SiC panels being filled by multiple layers of internal multiscreen insulation (IMI) consisting of a sandwich of platinum and gold sheets separated by ceramic spacers.

Propulsion:

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• Name: SDI RM 2100
• Type: Liquid-fuel engine
• Propellant: LOX/LH2
• Mixture Ratio: 7:1
• Cycle: Full-flow staged combustion
• Length: 4,300 mm
• Diameter: 2,400 mm
• Dry weight: 2,505 kg
• Nozzle expansion ratio: 60:1
• Thrust (vac):2,130 kN
• Thrust (sl): 2,050 kN
• Throttle Range: 100% to 20%
• Thrust-to-weight ratio: 83:1
• Chamber Pressure: 22.4 MPa
• Isp (vac): 450 seconds
• Isp (sl): 385 seconds

The first stage of the Taranis is powered by a single SDI RM 2100 rocket engine, a reusable liquid fuel engine w fueled by liquid hydrogen fuel and subcooled liquid oxygen oxidizer which is designed to fly 100 times between major maintenance intervals with a total life of 200 missions. The RM 2100 employs a full-flow, staged combustion (FFSC) cycle with a fuel rich preburner driving the high pressure fuel turbopump and an oxidizer-rich preburner driving the high pressure oxidizer turbopump. Liquid hydrogen fuel from the vehicle's fuel tank first passes through a fuel jet pump and then into the high pressure fuel turbopump where it then used to cool the main combustion chamber and nozzle. The vaporized fuel is then injected into the two preburners, the majority being infected into the fuel rich preburner and the rest into the oxidizer-rich preburner. Liquid oxygen oxidizer from the first stage LOX tank is first passed through an zidizer ejt pump and then into the high pressure oxidizer turbopump where the pressurized liquid oxygen oxidizer then passed through a heat exchanger where it becomes vaporized. A small amount of vaporized liquid oxygen oxidizer is used for pressurizing the main LOX tank while the rest is injected into the two preburners, the majority into the oxidizer rich preburner and the rest into the fuel rich preburner.