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Is That A Massive Stripe Streaking Across Pluto's Surface?

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The latest image of Pluto taken by the New Horizons spacecraft may have yielded the dwarf planet’s first prominent surface feature — a dark diagonal stripe that stretches from one side to the other.

New Horizons took these two images 15 seconds apart on June 6 at a distance of 45.8 million kilometers. The images have been embiggened by a factor of four.

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Here’s what Bjorn Jonsson of Unmanned Spaceflight has to say about these two photos and the apparent feature (emphasis mine):

The left version is the stack without any processing so it should show correct relative brightness. The right version has been sharpened using RegiStax. The sharpened version reveals a diagonal dark band on Pluto - it’s now absolutely certain that this is a real feature. In contrast, the apparently brighter terrain at the right limb is almost certainly a processing artifact. Charon may be starting to show large scale markings, i.e. possibly very slightly darker terrain in its upper left ‘quadrant’. But this could easily be an image processing artifact. [emphasis added] [...]

It wouldn’t surprise me if small dark spots started appearing within the bright terrain at much higher resolution and/or small bright spots started appearing within the dark terrain.

Over at The Planetary Society, Emily Lakdawalla quips: “Yay! Dark lines criscrossing a disk! It’s the discovery of canali on Pluto! We have reached Schiaparelli-quality mapping of Pluto’s surface!” She’s kidding of course...or is she?

...

:whistle:Nothing to see here:

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Delta IV Heavy

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The Delta IV Heavy is one of various versions of the workhorse launcher family of Delta IV Evolved Expendable Launch Vehicles. In the family of Delta IV rockets, the DIV Heavy is the most powerful launcher and, since the retirement of the Space Shuttle has been the most powerful launcher available in the United States, a role it will maintain until SpaceX introduces its Falcon Heavy booster.

The Delta family has a storied history with the original Delta making its first launch in 1960. Over the years, several generations of Delta launch vehicles have been flown - currently, the Delta II and IV rocket families remain in operation. Delta IV is built by Boeing, operated by United Launch Alliance for commercial launches, military missions as well as NASA Flights.

Delta IV rockets are operated from Space Launch Complex 37 at Cape Canaveral Air Force Station, Florida, and Space Launch Complex 6 at Vandenberg Air Force Base, California.

As one of the most powerful launch vehicles in the world, Delta IV heavy has a launch mass of 733 metric tons and can deliver payloads of 23 metric tons into orbit. The vehicle is comprised of a Common Booster Core as a central core stage with two additional CBCs attached to the core to deliver extra thrust in the early portion of the flight. Not employing propellant cross-feed, the two boosters fire at full throttle during their burn while the core saves propellants by throttling back shortly after launch to continue burning for another 90 seconds after strap-on separation.

Introduced in 2013, all Delta IV vehicles will transition to a more powerful version of the RS-68 cryogenic first stage engine known as RS-68A as soon as all existing RS-68 engines have been flown.

The Delta IV Heavy features the larger Delta Upper Stage with a 5-meter diameter and larger propellant tanks. The vehicle is outfitted with 5-meter payload fairings to accommodate large payloads. Delta IV is capable of delivering payloads to a variety of orbits such as Low Earth Orbit, Geostationary Transfer Orbit and Geostationary Orbit as well as interplanetary trajectories.

The entire Delta family makes use of flight proven components and proven design concepts to improve Launcher Success Rate. Delta IV Heavy has made seven launches to date (Nov. 2014). The maiden voyage of Delta IV Heavy in December 2004 was cataloged as a partial failure as all three CBCs suffered a premature cutoff during the launch sequence. The remaining launches were successful.

Other Delta IV Configurations: Delta Medium, M+ (4,2), M+(5,4), M+(5,2)

Specifications

Type Delta IV Heavy
Height 70.7m
Diameter 5m
Span 15.00m
Launch Mass 733,400kg
Stage 1 Common Booster Core
Boosters 2 Commom Booster Cores
Stage 2 5-Meter DCSS
Mass to LEO 22,950kg
Mass to GTO 13,130kg
Mass to GEO 6,275kg
Escape Capability 9,306kg

Launch Vehicle Description

The Delta IV Heavy Version consists of a Common Core Booster with two strap-on CBCs functioning as boosters attached to it, a 5-meter second stage and a 5-meter Payload Fairing. Each CBC uses a single RS-68 engine consuming cryogenic propellants, Liquid Oxygen and Liquid Hydrogen. Using the larger version of the Delta Cryogenic Second Stage, Delta IV Heavy sports a single RL-10B engine on its upper stage.

The Launcher stands 70.7 meters tall, has a main diameter of 5 meters and a liftoff mass of 733,400 Kilograms. With its three CBCs, the rocket has a span at its base of over 15 meters.

Delta IV Heavy rockets using the conventional RS-68 engines can lift payloads of up to 22,950 Kilograms to Low Earth Orbit. Geostationary Transfer Capability is 13,130 Kilograms, the vehicle can lift 6,275 directly to Geostationary Orbit and send payloads of up to 9,306 Kilograms to interplanetary trajectories. A typical Delta IV Mission has a duration of 2.3 hours, but can be extended to 7 hours for specific mission profiles.

Common Booster Core

The Delta IV Heavy uses a Common Booster Core as its first stage. This booster, like its name says, is common across all variants of the Delta IV launcher family. The Common Booster Core consists of a Main Engine Compartment, facilitating the RS-68 powerplant, a Liquid Hydrogen Tank, an Intertank Section, a Liquid Oxygen Tank and an interstage section that builds the interface with the Delta Cryogenic Upper Stage.

Both propellant tanks use an aluminum-isogrid structure with five isogrid structures welded together to make up the cylindrical section of the tanks. At both ends, the cylindrical tank section is welded to a bulkhead structure. Delta IV's CBC employs separate spherical bulkheads on the LH2 and LOX tanks. The tanks employ internal stringers for additional stability and the LOX tanks use anti-slosh baffles. The load-carrying external shell of the Common Booster Core is covered in rigid spray-on polyurethane foam for insulation and to prevent ice-build up around the cold cryogenic tanks. An external wiring tunnel runs down the length of the entire booster.

Sitting atop the LH2 tank, the 9.4-meter long Liquid Oxygen tank of the CBC holds 172,750 Kilograms (151 cubic meters) of the -183°C oxidizer at liftoff. LOX is fed to the RS-68 engine through a feedline that runs down the side of the CBC, external of the fuel tank.

The Liquid Hydrogen Tank is about 26.3 meters in length and capable of holding 29,500 Kilograms (416 cubic meters) of the -253°C cold substance directly fed to the engine via a short fuel supply line. Both, LOX and LH2 tanks are loaded with commodities via interfaces at the base of the launch vehicle.

The Common Booster Core is powered by a single RS-68 engine that, when flying in its RS-68A variant, is the most powerful Hydrogen-fueled engine in the world. The engine was developed by Rocketdyne Propulsion and Power, California and is now marketed by Aerojet Rocketdyne. RS-68 shares commonality with the Space Shuttle's RS-25 engines, but is overall of a simpler design to reduce manufacturing complexity and cost from $50 million for a single SSME to $14 million for RS-68. The engine requires about 80% fewer parts than the Shuttle's engine translating to a cut in labor of 92%.

Differences between RS-68 and RS-25 become evident in their turbopump design. While SSME required 170 individual parts for its LOX pump and 200 components were part of the LH2 pump, RS-68 only needs 40 LH2 pump parts and 25 parts for the LOX pump.

RS-68 provides 2,950 Kilonewtons of thrust at liftoff and has a vacuum thrust of 3,370 Kilonewtons. It has a dry weight of 6,747 Kilograms and provides a specific impulse of 409 seconds. The engine can be throttled between 55% and 100% of rated thrust using a simple single-step throttle profile. It operates at a propellant mixture ratio of 5.97 and has an expansion ratio of 21.5, operating at a chamber pressure of 196 bar.

The RS-68 is a cryogenic booster engine using a basic gas generator cycle. The engine consists of a gas generator driving the turbines of two independent turbopumps feeding a combustion chamber that is using a channel-wall design. Inner and outer skins brazed to middle separators form cooling channels as part of a simple engine cooling design using the flow of hydrogen to cool the chamber while the engine nozzle uses ablative cooling through an ablative layer that slowly burns away during operation of the engine.

Employing an open-cycle design, the RS-68 consists of a central Gas Generator that uses part of the fuel and oxidizer flow from the two turbopumps of the engine that are spun-up by pressurized helium during engine start. To start up, RS-68 first opens its Gas Generator and Main Fuel Valves and GG LOX valve at T-5.5 seconds, applying helium for turbopump sin-up and actuating engine igniters. This ensures a fuel-rich engine start needed to maintain safe engine temperatures during the ignition sequence which can never be oxidizer-rich. The LOX side opens up its Main Valve at T-2.0 seconds and spins up its turbopump, delivering high-pressure oxidizer to the Gas Generator to sustain the operation of the turbines of the LOX and LH2 turbopumps to deliver propellants to the engine.

The high-pressure gas from the Gas Generator is divided into two outflow paths, one over to the LOX pump's turbine, the other over to the fuel-turbopump. After driving the turbines, the gas is dumped overboard. On the oxidizer side, the gas generator gas flows through a heat exchanger that uses a portion of the LOX flow through the Main Oxidizer Valve to generate gaseous oxygen that is used for tank pressurization - delivered to the LOX tank through a pressurization line that includes pressure regulators to control the tank pressure.

The turbine gas from the LH2 side is directed through a movable nozzle that is actuated for roll control of the Common Booster Core when flying without its strap-on CBCs later in the mission. After passing through the Main Fuel Valve, the Liquid Hydrogen is run through the regenerative cooling cycle of the combustion chamber. Transitioned to a gaseous state, part of the Hydrogen flow is used to pressurize the LH2 tank during flight. Pre-launch pressurization of the propellant tanks is accomplished using Helium gas.

Launch Vehicle Control during first stage flight is provided by the main engine. Pitch and Yaw are controlled by gimbaling the engine while Roll Control is accomplished by vectoring the RS-68 turbine exhaust gases.

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First Stage

Type Common Booster Core
Inert Mass 26,400kg
Launch Mass 228,400kg
Diameter 5.1m
Length 40.8m
Propellant Liquid Hydrogen
Oxidizer Liquid Oxygen
Fuel&Oxidizer Mass 202,000kg
Tank Structure Al-Isogrid, Separate Bulkheads
LOX Tank Length 9.4m
LH2 Tank Length 26.3m
LOX Mass / Volume 172,500kg / 151m³
LH2 Mass / Volume 29,500kg / 416m³
Tank Pressurization Gasified Propellants
Guidance From 2nd stage
Propulsion 1 RS-68
Cycle Open Cycle, Gas Generator
Thrust (SL) 2,950kN
Thrust (Vacuum) 3,370kN
Chamber Pressure 196bar (at 100% Throttle)
Engine Length 5.20m
Engine Diameter 2.43m
Engine Dry Weight 6,747kg
ISP (Sea Level) 359s
ISP (Vacuum) 409s
Mixture Ratio 5.97
Nozzle Ratio 21.5
Throttle Capability 55%-102%
Pitch, Yaw Control Engine Thrust Vector Control
Roll Control Vectoring of GG Exhaust Nozzle
Restart Capability No
Burn Time 328sec
Stage Separation Pyro bolts, Springs

Strap-On Common Booster Cores

Delta IV Heavy uses two Common Booster Cores that are attached to the CBC in the center in a strap-on fashion to serve as liquid-fueled boosters, delivering additional thrust for the first minutes of flight. The CBCs functioning as boosters are attached to the central core using thrust struts that interface with the interstage section of the launcher to transfer loads from the boosters to the rest of the vehicle. Additional attachment points reside in the base of the vehicle right above the engine heat shields. All interface points include pyrotechnic separation devices that are used to jettison the boosters after burnout.

Delta IV Heavy does not have a propellant crossfeed capability. For launch, all three CBCs are fired at full thrust before the RS-68 engine of the central core throttles down to 55% of rated performance to save propellants. Depending on mission requirements, the throttle-down of the core occurs approximately 50 seconds into the flight.

The two Common Booster Cores functioning as boosters continue flight at full thrust, consuming all their propellants in 242 seconds with separation coming two seconds after RS-68 cutoff.

By running on 55% for the initial portion of the flight, the Common Booster Core has propellants left to continue to power the vehicle after Booster Jettison. The RS-68 Engine throttles back up to 100% and burns for 88 seconds following booster separation. At T+328 Seconds, the Common Booster Core shuts down and separates from the vehicle shortly thereafter using pyrotechnic bolts and springs that push the spent CBC away.

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Strap-On Booster

Type Common Booster Core
Inert Mass 26,400kg
Launch Mass 228,400kg
Diameter 5.1m
Length 40.8m
Propellant Liquid Hydrogen
Oxidizer Liquid Oxygen
Fuel&Oxidizer Mass 202,000kg
Tank Structure Al-Isogrid, Separate Bulkheads
LOX Tank Length 9.4m
LH2 Tank Length 26.3m
LOX Mass / Volume 172,500kg / 151m³
LH2 Mass / Volume 29,500kg / 416m³
Tank Pressurization Gasified Propellants
Guidance From 2nd stage
Propulsion 1 RS-68
Cycle Open Cycle, Gas Generator
Thrust (SL) 2,950kN
Thrust (Vacuum) 3,370kN
Chamber Pressure 196bar (at 100% Throttle)
Engine Length 5.20m
Engine Diameter 2.43m
Engine Dry Weight 6,747kg
ISP (Sea Level) 359s
ISP (Vacuum) 409s
Mixture Ratio 5.97
Nozzle Ratio 21.5
Throttle Capability 55%-102%
Pitch, Yaw Control Engine Thrust Vector Control
Roll Control Vectoring of GG Exhaust Nozzle
Restart Capability No
Burn Time 242sec
Stage Separation Pyro bolts cutting Thrust Struts

Second Stage

The Delta IV Heavy launch vehicle uses the larger of the Delta upper stages that has an increased diameter of five meters. The Delta Cryogenic Second Stage with a five-meter diameter has a stretched version of the original LOX tank of the four-meter stage, still remaining at the original diameter of 3.2 meters while the LH2 tank has its diameter widened to five meters, allowing the stage to carry a total of 27,220 Kilograms of cryogenic propellants for consumption by the RL-10B engine of the Upper Stage.

The LH2 tank is located above the LOX tank with a truss structure connecting the two tanks that use individual bulkheads. The space in between the two tanks is used to facilitate helium pressurization bottles, attitude control system tanks and the launch vehicle’s avionics known as Redundant Inertial Flight Control Assembly (RIFCA), located on a shelf below the LOX tank.

Overall, the LOX tank of the DCSS measures 4.0 meters in length while the LH2 tank is 4.8 meters long. The entire engine compartment and LOX tank as well as the intertank section and lower segment of the LH2 tank are protected in the interstage section that remains with the Common Booster Core after separation. The upper section of the LH2 tank and the payload adapter are protected by the payload fairing so that only the 5-meter segment of the LH2 tank is actually visible on the exterior of the launcher when in its liftoff configuration.

The Delta Cryogenic Second Stage is powered by a single RL-10B-2 engine that is part of the RL-10 family that can look back on a history of several decades having completed its first test in 1959 after being developed by Pratt & Whitney. Over the years, the engine underwent a number of modifications, going through several generations used on different launch vehicles. The RL-10A version of the engine has been powering the trusted Centaur upper stage for decades, currently used in its RL-10A-4-2 evolutionary stage while previous RL-10A version were employed by the Saturn I and DC-X vehicles.

The RL-10B-2 version of the engine with enhanced performance through an extended nozzle is only used by the Delta IV launcher. Another version of the engine, RL-10C is currently being tested for use on Centaur and a derivative of the engine known as Common Extensible Cryogenic Engine (CECE) has completed demonstration tests that included a record-setting deep throttling capability.

The RL-10B engine delivers 110 Kilonewtons of thrust at a specific impulse of 464 seconds employing an expander cycle. The engine features an extendable nozzle that utilizes an electromechanical system that moves the radiatively cooled nozzle extension into position on the regeneratively cooled nozzle segment immediately after stage separation. Overall, the engine is 4.14 meters long with its extension in place, creating a diameter of 2.21 meters and a high expansion ratio of 250, optimized for operation in vacuum conditions. The engine has a dry mass of 277 Kilograms and operates at a mixture ratio of 5.88.

RL-10 is a closed Expander Cycle Engine which does not rely on a gas generator to deliver the hot gas that drives the turbopump turbines of the engine. Instead, the turbines are driven by expanded hydrogen gas that is generated by running the flow of Liquid Hydrogen from the LH2 turbopump through the regenerative cooling system of the upper nozzle segment and the combustion chamber. The gasified Hydrogen then passes to the main turbine of the engine, spinning the LH2 turbopump as well as the LOX turbopump through a gearbox.

Rl-10 includes seven engine valves starting on the fuel side with the Fuel Pump Inlet Shutoff Valve and on the oxidizer side with the Oxidizer Pump Inlet Shutoff Valve. Fuel flow into the combustion chamber can be stopped by the Fuel Shutoff Valve that is located just upstream of the combustion chamber injector. This valve is used to rapidly cut the fuel feed to the engine for shutdown and its closure also allows the chilldown of the LH2 turbopump through overboard vents without any fuel entering the chamber. Engine LH2 pump chilldown is accomplished by opening Fuel-Cool-Down Valves 1 & 2 that vent coolant overboard during chilldown. These two valves also provide fuel pump bleed during pre-start and pressure relief during shutdown.

Thrust of the engine is controlled by a Thrust Control Valve located in a bridge between the fuel cooling outlet on the engine and the combustion chamber fuel inlet to bypass the turbine and thus regulate turbine power and overall engine thrust. Normally in a closed position, the system is mainly used to control thrust overshoot during engine start and to maintain a constant chamber pressure during steady state operation.

In the oxidizer line downstream of the pump is a Oxidizer Flow Control Valve that is used to regulate the LOX flow to the chamber for the regulation of the mixture ratio that is commanded by the Propellant Utilization Unit of the engine which controls the MR for an optimized propellant consumption. A second OCV is employed to regulate the bleed flow during engine start.

Engine start on the RL-10 is accomplished by using the pressure differential between the fuel feed and the near-vacuum in the chamber that forces fuel through the system after the Fuel Shutoff Valve is opened and FCV-1 is closed. FCV-2 remains in an open position to prevent stalling the LH2 pump of the engine in start-up. In the initial stages of start-up, heat from the ambient metal is sufficient to generate Hydrogen gas to start driving the turbopumps and initiate the combustion process in the chamber, heating up the chamber and nozzle to operational levels. For start, the Oxidizer Control Valve is partially closed to ensure a fuel-rich ignition, limiting chamber pressure in order to maintain a pressure differential in the fuel system until the turbopumps can accelerate.

When the pumps are at flight speed, pneumatic pressure is used to close the Fuel Cooldown Valve and open the Oxidizer Control Valve to achieve the planned LOX pump discharge properties. The opening of the OCV leads to a sharp increase in chamber pressure that can lead to thrust overshoot which is prevented by a temporary opening of the Thrust Control Valve until stable steady-state conditions are reached for engine operation.

In steady state operation, RL-10 consumes 20.6 Kilograms of LOX per second while LH2 flow is approximately 3.5 Kilograms per second.

Engine shutdown is a simple process accomplished by closing the Fuel Shutoff and Fuel Inlet Valve and at the same time opening the Fuel Control Valves to bleed fuel from the system. Oxidizer flow is cut by closing the Oxidizer Control Valve and LOX Inlet Valve. Friction losses lead to the spin-down of the turbines and pumps.

The DCSS uses high-pressure Helium to keep its LOX tank at flight pressure while the LH2 tank uses gaseous hydrogen from the engine bleed that is delivered via regulators that ensure proper tank pressurization.

Propellant management is accomplished by directing hydrogen boil-off from the tank to aft-facing thrusters that deliver sufficient thrust for propellant settling, keeping up a uniform two-phase system between liquids and gases within the propellant tanks. Propellant settling can also be provided by the attitude control system of the second stage.

The attitude control system of the Delta Cryogenic Second Stage consists of 12 hydrazine monopropellant thrusters installed in four modules on the upper stage to deliver control around all three axes as well as additional propellant settling capability. Each module features three thrusters, two lateral and one axial. Fed from several hydrazine bottles, the thrusters use the catalytic decomposition of hydrazine over a metallic catalyst bed.

The MR-106D thrusters are manufactured by Aerojet Rocketdyne with each lateral thruster delivering 21 to 41 Newtons of thrust (smaller 17-27N versions of the MR-106D are not used by DCSS). MR-106D can operate at propellant pressures of 13.8 to 21 bar with chamber pressures varying from 8.6 to 17.2 bar. Flow rate varies between 9.5-17.7g/sec. The thrusters use single seat valves operating at 28 Volts. The engines create a specific impulse of 227 to 234 seconds and are capable of operating in pulse mode with a minimum impulse bit of 0.63Ns while operation in steady state mode up to 1,000 seconds is also possible. Each thruster weighs about 0.6kg with an overall mass per Rocket Engine Module of 2.5 Kilograms.

DCSS control in yaw and pitch is provided by an electromechanical Thrust Vector Control System that gimbals the RL-10 main engine while roll control relies on the attitude control system. The ACS is responsible for three-axis control during coast phases including thermal control maneuvers and re-orientations for spacecraft separation.

The DCSS also facilitates RIFCA, the Redundant Inertial Flight Control Assembly which provides primary guidance and control of all stages of the Delta IV flight. Manufactured by L-3 Space & Navigation, the RIFCA made its debut in 1994. The unit is triple redundant and measures 43 by 36 by 24 centimeters in size with a mass of 33 Kilograms and a power demand of 75 Watts. The system is connected to all launch vehicle actuators and sensors using a MIL-STD-1553 data bus. RIFCA uses six ring laser gyros and six accelerometers to accurately determine the vehicle’s attitude and body rates operating at rates of up to 30 degrees per second and achieving an accuracy of under 0.08 degrees.

Powered by batteries, the DCSS and operate for 2.3 hours in its standard configuration, although batteries can be added to support missions up to 7 hours that also require additional hydrazine tanks for attitude control system propellant. After payload separation, the second stage can stay active to make a Contamination and Collision Avoidance Maneuver or a Deorbit Burn before passivation.

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Second Stage

Type Delta Cryogenic Upper Stage
Diameter 5m
Length 13.7m
Inert Mass 3,490kg
Propellant Liquid Hydrogen
Oxidizer Liquid Oxygen
Fuel&Oxidizer Mass 27,220kg
Tank Structure Al-Isogrid, Separate Bulkheads
LOX Tank Length 4.0m
LH2 Tank Length 4.8m
Tank Pressurization Gasified Propellants
Guidance Inertial from RIFCA
Propulsion 1 RL-10B-2
Thrust 110kN
Engine Length 4.14m
Engine Diameter 2.21m
Engine Dry Weight 277kg
Specific Impulse 464sec
Nozzle Ratio 250:1
Nozzle Extension Yes
Ox. To Fuel Ratio 5.88:1
LOX Flowrate 20.6kg/sec
LH2 Flowrate 3.5kg/sec
Burn Time Variable (1,125sec)
Engine Start Spark Igniter, Restartable
Attitude Control RL-10 Gimbaling (Pitch, Yaw)
ACS Redundant Attitude Control System
ACS Propellant Hydrazine
ACS Thrusters 12 x MR-106D (4 Pods)
MR-106D Thrust 21 - 41 N
Specific Impulse 227 - 234 s
Chamber Pressure 8.6 - 17.2 bar
Propellant Pressure 13.8 - 21.0 bar
Flowrate 9.5 - 17.7 g/sec
Min. Impulse Bit 0.63Ns
Max Burntime 1,000sec
Thruster Mass 0.6kg

Payload Adapters/Fittings

Payload Adapters interface with the vehicle and the payload and are the only attachment point of the payload on the Launcher. They house equipment that is needed for Spacecraft Separation and ensure that the satellite or spacecraft is secured during powered flight.

A variety of payload adapters is available to satellite customers in order to fit a large number of spacecraft dimensions and interfaces. 7 different adapter models are currently available. Most of those have a clamp band payload separation mechanism. It is also possible to design new adapters to accommodate a variety of spacecraft.

Payload Fairing

The Payload Fairing is positioned on top of the stacked vehicle and its integrated payloads. It protects satellites or other spacecraft against aerodynamic, thermal and acoustic environments that the vehicle experiences during atmospheric flight. When the launcher has left the atmosphere, the fairing is jettisoned by pyrotechnical initiated systems. Separating the fairing as early as possible increases ascent performance. Payload Fairing design limits Payload Volume.

For the Delta IV Heavy Version, only 5-meter diameter fairings can be used, because smaller fairings do not fit on top of the large second stage. Two different versions with lengths of 19.1 and 19.8 meters are available to accommodate different payload dimensions. The smaller version of the DIV heavy Fairing uses the conventional Graphite-Epoxy Sandwich Structure whilst the large 19.8-meter is an Aluminum Isogrid Structure.

The fairing is separated by pyro bolts and spring jettison actuators that push the two halves away from each other. Payload Fairings are outfitted with acoustic panels, access doors and RF windows. Also, the Payload Fairing is connected to a purge air system to ensure a controlled environment. Off-pad encapsulation is provided to improve safety standards and limit on-pad time.

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Payload Fairing

Type Long Delta IV Fairing
Diameter 5m
Length 19.1m
Separation Pyrotechnic Activation (Actuators)
Construction Sandwich Construction
Graphite-Epoxy/Foam Core

Type Long Delta IV Fairing
Diameter 5m
Length 19.1m
Separation Pyrotechnic Activation (Actuators)

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Pegasus XL Launch Vehicle

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The Pegasus XL Launch Vehicle is a light-weight lift vehicle operated by Orbital Sciences Corporation. It can deliver Payloads of up to 443 Kilograms into Low Earth Orbit. The Rocket is launched from the L-1011Stargazer Aircraft which makes Pegasus a high flexibility launch system as it is not depending on a fixed Launch Location and associated ground weather. Also, launching from an Aircraft enables the vehicle to reach a variety of Orbits with different inclinations including nearly equatorial orbits. To date, Pegasus has made 40 Launches, 35 of which were successes and 2 partial failures. Pegasus made its first flight in 1990 delivering two satellites to orbit. The vehicle is a three-stage, solid propellant rocket with the option of a fourth, liquid fueled stage. The Launch System is a low cost option for smaller payloads not requiring precision injections.

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Vehicle&Flight Description

First Stage

At T-0 the Launch Vehicle is released by the L-1011 Aircraft and free-falls for five seconds before igniting its Orion 50S Solid Rocket Motor - the fist Stage of the Vehicle. It burns for about 69 seconds before shutting down at an altitude of about 61 Kilometers. During first stage flight, the Rocket is controlled by its fin actuators that provide pitch control. The Orion 50S does not have Thrust Vector Control Capabilities. A Delta Wing provides some lift and supports a pitch maneuver that is initiated shortly after ignition. After first stage cutoff, the vehicle holds on to the stage for several more seconds before stage separation occurs. The Wings and Fins are jettisoned with the first stage.

Second Stage


The second stage is similar to the first stage, however it uses thrust vector control and closed loop guidance to control the vehicle. Orion 50 also uses sold propellant and burn for nearly 70 seconds. Halfway through second stage flight, the payload fairing separates from the vehicle and fall into the ocean. Attitude control is provided by Thrust Vector Control for Pitch and Yaw and Nitrogen Thrusters for Roll Control.

Third Stage

After second stage cutoff and subsequent stage separation, the Orion 38 Solid Rocket Motor takes over powered flight and places the Payload in its desired orbit or injects the upper composite consisting of HAPS and Payload into a transfer orbit. Third Stage Navigation and Control is similar to that of the second stage. A typical Pegasus Ascent takes around 9.5 Minutes from Launch Vehicle Drop to Spacecraft Separation.

Fourth Stage

An optional fourth stage can be added to the Launch Vehicle stack to increase ascent performance or provide high-precision injections into a variety of orbits. The HAPS uses liquid propellants and provides a multi-ignition capability to enable payloads to be injected accurately. The fourth stage further limits payload volume as it has to fit inside the Payload Fairing with all associated interfaces and the actual Satellite.

Payload Fairing

The Payload Fairing is positioned on top of the stacked vehicle and its integrated payloads. It protects satellites or other spacecraft against aerodynamic, thermal and acoustic environments that the vehicle experiences during atmospheric flight. When the launcher has left the atmosphere, the fairing is jettisoned by pyrotechnically initiated systems. Separating the fairing as early as possible increases ascent performance. Payload Fairing design limits Payload Volume.

Payload Adapters

Payload Adapters interface with the vehicle and the payload and are the only attachment point of the payload on the Launcher. They house equipment that is needed for Spacecraft Separation and ensure that the satellite or spacecraft is secured during powered flight. Off-the-shelf and custom adapters are available to customers to accommodate a variety of payloads.

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Pegasus Specifications

Type Pegasus XL
Height 16.9m
Diameter 1.27m
Launch Mass 23,130kg
Stages 3
Stage 1 Orion 50S XL
Stage 2 Orion 50 XL
Stage 3 Orion 38
Stage 4 - Optional HAPS
Mass to LEO 443kg
Launch Cost ~$11 Million (1994)

First Stage

Type Orion 50S XL
Inert Mass 1,369kg
Diameter 1.28m
Length 10.27m
Propellant Solid
Propellant Mass 15,014kg
Guidance Open Loop
Propulsion Orion 50S XL
Thrust (Vacuum) 726kN
Impulse 295s
Average Pressure 1,090psia
Burn Time 68.6sec
Attitude Control Fin Actuators

Second Stage

Type Orion 50 XL
Diameter 1,28m
Length 3.11m
Inert Mass 416kg
Propellant Solid
Propellant Mass 3,925kg
Guidance Closed Loop PEG
Propulsion Orion 50 XL
Thrust 196kN
Impulse 289s
Burn Time 69.4sec
Average Pressure 1,019psia
Attitude control Thrust Vector Control
Reaction Control System (GN2)

Third Stage

Type Orion 38
Diameter 1.34m
Length 0.97m
Inert Mass 126kg
Propellant Solid
Propellant Mass 770kg
Guidance Closed Loop PEG
Propulsion Orion 38
Thrust 36kN
Impulse 287s
Burn Time 68.5sec
Average Pressure 656psia
Attitude control Thrust Vector Control
Reaction Control System (GN2)


Fourth Stage

Optional: Hydrazine Auxiliary Propulsion System

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’Repackage Our Margin’

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More than two decades have passed since one of the most spectacular EVA missions in U.S. history: the long-awaited first servicing of the Hubble Space Telescope (HST). Launched in April 1990, the $1.5 billion observatory was the jewel in NASA’s scientific crown, but shortly afterwards fell foul to the effects of a spherical aberration in its primary optics, which severely impaired the quality of its images. With funding for Space Station Freedom—later to evolve into the International Space Station (ISS)—hanging on the edge of a knife, a successful repair and rejuvenation of Hubble was critical not only to the telescope’s future, but to the future of NASA itself. In December 1993, the seven-member crew of Endeavour launched on STS-61, an 11-day flight to perform a record-setting series of five EVAs to bring Hubble back from the brink of disaster and restore the space agency’s tattered reputation. In doing so, they accomplished one of the most spectacular human space missions of the decade and demonstrated the shuttle’s capabilities for the future.

Central to the repair effort was the $50 million Corrective Optics Space Telescope Axial Replacement (COSTAR), fabricated by Ball Aerospace, which would correct the spherical aberration by positioning 10 small, coin-sized mirrors to restore the potential of Hubble’s affected scientific instruments. However, in order to make room for COSTAR, another instrument—the phone-booth-sized High Speed Photometer (HSP), rendered useless by “jittering” in the telescope’s solar arrays—would need to be removed and returned to Earth. “Once in place,” explained NASA’s STS-61 Press Kit, “COSTAR will deploy a set of mechanical arms, no longer than a human hand, that will place corrective mirrors in front of the openings that admit light” into the affected instruments. In doing so, it would refocus light from Hubble’s primary mirror before it reached those instruments and was expected to bring their overall optical performance “very close” to original specifications. Follow-on instruments for the telescope would be specifically designed with their own corrective optics already pre-integrated.

Additional tasks included the installation of a new Wide Field Planetary Camera (WFPC-2), the replacement of Hubble’s twin solar arrays and drive electronics, two of three Rate Sensing Units (RSUs), one of two Electronic Control Units (ECUs), one of two magnetometers, and fuse plugs to correct wiring discrepancies. In March 1992, Story Musgrave was assigned as payload commander for the flight, and in August fellow astronauts Jeff Hoffman, Kathy Thornton, and Tom Akers joined him to support at least five ambitious EVAs. The final members of the crew—Commander Dick Covey, Pilot Ken Bowersox, and European Space Agency (ESA) astronaut Claude Nicollier—were named in December 1992, tracking a December 1993 launch. With the repair work, STS-61 had morphed into a far more complex mission than had been anticipated before Hubble’s launch.

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In its January 1990 manifest, the agency listed SM-1 as a five-day flight with a crew of five, suggesting a maximum of only two or three EVAs, but as 1991 wore into 1992 and onward into 1993 it became increasingly clear that the mission would run to as long as 11 days and evaluations of underwater simulations convinced managers that they should schedule as many as five back-to-back EVAs over five days. According to Mission Director Randy Brinkley, the decision served to “repackage our margin” and offered the chance to “respond to the dynamics, or unknowns, of spacewalks.” (The flight plan actually provided for a sixth and seventh EVA, and a mission duration of up to 13 days, although this was did not become necessary.)

Such an enormous workload demanded a crew of seven, with two alternating teams of spacewalkers, to reduce fatigue and enhance the likelihood of mission success. Original plans called for all tools to be kept outside, in the shuttle’s payload bay, but the crew recognized at an early stage that EVA time was a critical limiting consumable and decided that the hour spent preparing equipment at the start of each excursion could be better spent starting the repair work. It was therefore decided that some tools would be kept inside Endeavour’s crew cabin, enabling the spacewalkers to “load-up” before opening the airlock and utilizing their suits’ consumables. “What we’ve done by going to five EVAs, rather than three, is to repackage our margin,” said Brinkley, “so that we have the capability to respond to the dynamics, or unknowns, of spacewalks. It improves the probabilities for mission success, while providing added flexibility and adaptability for reacting to real-time situations.”

“Doing five [EVAs] really pushed the bounds of what people thought we could do,” Covey recalled in his NASA oral history. “Even with four EVA crew members, even with an 11-day mission, it just started pushing the bounds. There was a lot of scrutiny on it and a lot of focus on it.” The size of their quarry posed additional problems. Hubble was far larger than anything with which the shuttle had previously rendezvoused in orbit, and Claude Nicollier was faced with the unenviable challenge of maneuvering his EVA crewmates, along with phone-booth-sized pieces of hardware, into position by means of the Remote Manipulator System (RMS) mechanical arm with extreme delicacy and precision. “The integrated operations,” said Covey, “of shuttle maneuvering, RMS activities and EVAs, although now commonplace, wasn’t back then. So integrating all of those activities and the crew activities together was a big part of my role as the commander.”

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All of the spacewalkers recognized the need to develop physical strength to handle the demands of their space suits and build the necessary stamina for six or seven hours outside. Kathy Thornton worked out in the gym, as did the others, although by Hoffman’s admission most of the servicing tasks did not demand immense physical strength, but placed greater emphasis on “technical co-ordination,” involving them “being very careful in how you moved things around and not messing anything up.”

Following a successful launch on 2 December 1993, it was recognized that the spacewalks would be performed daily, with Musgrave and Hoffman charged with the first, third, and fifth and Thornton and Akers assigned to the second and fourth. Encased within their pressurized suits, the astronauts were identified by the presence (or absence) of markings on their legs: Hoffman (EV1) would have red stripes, Musgrave (EV2) would have no stripes, Thornton (EV3) would have dashed red stripes, and Akers (EV4) would have diagonal broken red stripes. All four spacewalkers were extensively “cross-trained” to allow them to perform any one of the mission’s given EVA tasks and around 200 tools, from power ratchets and sockets to safety bars and articulating foot restraints and from portable work lights and locking connectors to instrument covers, handles, and umbilical connectors.

Looking back on those adrenaline-charged days, Covey was filled with pride that his crew accomplished everything they set out to do. “There wasn’t anybody that was chilling down on the middeck,” he said. “Everybody was up top, working. There was concern about whether we could sustain that tempo. We went five straight days doing EVAs and that was the right answer. Everybody felt good about that. Nobody was getting excessively fatigued. The EVA crew members, because they were getting a day off in between were okay with that and so that facilitated us pressing on with five straight days of spacewalks.”

All five EVAs were significant in reviving Hubble, but in the eyes of the public and politicians perhaps the most critical tasks were the installation of WFPC-2 by Musgrave and Hoffman on EVA-3 on 6/7 December 1993 and the installation of COSTAR by Thornton and Akers on 7/8 December. Before launch, Hoffman remembered being told not to worry if they did not accomplish everything on the manifest; as long as eitherWFPC-2 or COSTAR was successfully installed, the scientists on the ground would be “deliriously happy.” However, they were not fully appreciative of NASA’s collective mindset of having a 100-percent-successful mission.

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WFPC-2 had originally been developed in 1985 as a “spare,” but after the discovery of the spherical aberration NASA had requested the installation of an optical corrector. “The new design incorporates an optical correction by the refiguring of relay mirrors already in the optical train of the cameras,” read NASA’s pre-flight press kit. “Each relay mirror is polished to a new specification that will compensate for the incorrect figure on [Hubble’s] primary mirror. Small actuators will fine-tune the positioning of these mirrors on-orbit, ensuring the very precise alignment that is required.” The WFPC team also upgraded the instrument, by reducing the number of cameras from eight to four in order to develop an alignment system and adding improved charge-coupled devices to aid its ultraviolet sensitivity.

An hour into the spacewalk, Hoffman crisply removed the original WFPC-1 from its housing in Hubble’s bowels and inserted it into a storage container in the payload bay. A protective hood was then removed from the new device and it was installed perfectly. Ground controllers ran an “aliveness” test and verified that the pie-wedge-shaped WFPC-2 was working correctly. The spacewalkers then replaced a pair of magnetometers, before returning inside Endeavour after six hours and 47 minutes. This proved exceptionally good time, when one considers that training for the WFPC-2 replacement alone had typically taken 4.5 hours in the water tank.

A day later, Thornton and Akers set about exchanging the HSP for COSTAR. This required them to open the telescope’s bay doors and loosening latches and removing electrical connectors in order to slide out the instrument. The new corrective optics package was then fitted. In training on Earth, the operation had taken around 3.5 hours. The intensity of the mission—an intensity which had impacted Story Musgrave for almost two years, to such an extent that he remarked, with the merest hint of jest, that the only peace and solace he could find from the mission was sitting in the dentist’s chair—began to lessen somewhat when Thornton and Akers successfully removed the photometer and installed COSTAR in its place. By the end of the EVA, both WFPC-2 the corrective optics had been triumphantly fitted.

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Although these two EVAs restored much of Hubble’s science-gathering capability, the other three excursions of STS-61 replaced its solar arrays and other critical equipment and transformed the telescope into virtually a new spacecraft. From an EVA perspective, the records fell like ninepins during STS-61. By the time EVA-5 was completed, the four spacewalkers—Musgrave, Hoffman, Thornton, and Akers—had totaled more than 35 hours outside Endeavour and five excursions on a single flight was more than had ever been achieved on a single shuttle mission. Moreover, STS-61 was the first shuttle flight in which the bounds of accomplishment, in terms of mission duration, complexity, and integrated EVA-RMS-orbiter operations, were pushed to their absolute limits.

That night, the night after the final EVA, the crew of STS-61 celebrated their success above the roof of the world. “Of all of the programs that I have been associated with,” Dick Covey remembered, years later, “it’s the one that was best planned and has been best executed, in terms of using astronauts and crewed vehicles to be able to support, enable and enhance the scientific mission of space.” STS-61 had done nothing less than save NASA itself. Few other human space missions since Apollo 11 had exerted such a positive influence on the agency’s subsequent fortunes. Of course, we know today that fixing Hubble’s optics was triumphantly successful, and the telescope repair team received the prestigious Robert J. Collier Trophy in May 1994 for their work. The citation praised their “outstanding leadership, intrepidity and the renewal of public faith in America’s space program by the successful orbital recovery and repair of the Hubble Space Telescope.”

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Saturn's Massive Phoebe Ring is Even Larger Than Previously Thought

Saturn’s Massive Phoebe Ring is Even Larger Than Previously Thought « AmericaSpace

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Saturn is truly the “Lord of the Rings” and one of the most majestic places in the Solar System. It’s massive ring system is well-known, but in 2009 another previously unknown ring was discovered, much larger than the others but fainter, being composed of dark grains of dust thought to originate from the moon Phoebe. Now, new research indicates that the Phoebe ring is even larger than first thought.

The new findings come from NASA’s Wide-field Infrared Survey Explorer (WISE) spacecraft; the tenuous ring is now seen to extend from 3.75 million to 10.1 million miles (6 million to 16.2 million km) from the planet, or about 100 to 270 times the radius of Saturn itself. Previous estimates from NASA’s Spitzer Space Telescope said the ring extended distances of 128 to 207 times the radius of Saturn, or about 4.8 million to 7.7 million miles (7.7 million to 12.4 million kilometers). Even then, that would be about 12.5 times the average distance between Earth and the Moon, more than 10 times larger than Saturn’s next largest and more visible ring, the E ring. This massive, diffuse torus-like ring dwarfs Saturn’s other rings and the planet itself.

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As noted by Douglas Hamilton, a planetary scientist at the University of Maryland, College Park, inSpace.com, “We knew it was the biggest ring, but know we find it’s even bigger than we thought, new and improved.”

As he told NPR, the ring is “more than 200 times as big across as Saturn itself – it’s absolutely immense, much bigger than any other ring that we know of.”

“It’s fascinating that this ring can exist,” Hamilton continued. “We’re told in science textbooks that planetary rings are small and close to their parent planets – if they’re too far away from their planets, moons form rather than rings. This discovery just turns that idea on its head – the universe is a more interesting and surprising place than we thought.”

The findings were just published in the June 11 issue of Nature.

Another moon, Iapetus, had also provided clues to the existence of the new ring.

“Like our moon, Iapetus always has one side facing toward Saturn, which means it also always has one side pointing in the direction of its motion around Saturn, its leading side,” Hamilton said. “Iapetus is an icy moon, and intrinsically bright white, but its leading face is very strikingly jet black. That contamination is what led us to look for what turned out to be a surprisingly large ring.”

The dust particles in the ring are exceedingly tiny, only about 10 to 20 microns in size. They make up the bulk of the ring, with larger particles, up to about 7.8 inches (20 centimeters) or more, making up no more than 10 percent of the ring material. The Phoebe ring is darker and less prominent due to the dust particles being darker themselves and most likely coming off of the dark moon Phoebe. The particles absorb sunlight so the ring is less distinct in visible light, but more visible in infrared light. This makes it easier to see the ring in the infrared images taken by WISE. The particles are also well spread out, with most of the ring being empty space, kind of like atoms.

“A cubic kilometer of space in the Phoebe ring might have just a few dozen dust particles, maybe 100 at most,” Hamilton said. “It’s really empty.”

The ring has also been seen in optical light however, as reported in 2014, when it was observed by the ISS camera on the Cassini spacecraft.

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The particles are also thought to be millions to billions of years old, since there is little chance of any of them colliding and destroying each other sooner in collisions. The findings also show how much variety there actually is in Saturn’s rings.

As Hamilton summarized, “Saturn’s main rings are like the fabled elephant graveyard – mysterious and filled with mostly large bones that contain clues about the recent past,” Hamilton said. “The E ring, then, is the chipmunk graveyard in which all of the bones are small and from the modern era, and the Phoebe ring is the dinosaur graveyard in which we find ancient bones of all sizes, most of them tiny fragments but some quite immense.”

Hamilton and his colleagues also think it is possible that Jupiter has a similar, as-yet undiscovered larger ring (besides its other already known rings which are visible but fainter than Saturn’s).

“Whenever a planet has a distant satellite, it will probably have a distant ring as well. We see Saturn’s because it’s bright enough to image; Jupiter’s is probably fainter and harder to spot.”

In related news, the Cassini spacecraft has taken beautiful new images of the moon Tethys, froma distance of approximately 118,000 miles (190,000 kilometers) and with an image scale of 3,280 feet (1 kilometer) per pixel. The largest crater, Odysseus, is 280 miles (450 kilometers) across, covering about 18 percent of the moon’s surface area; Tethys itself is only 660 miles (1,062 kilometers) across. For comparison, a similarly sized crater on Earth would be as big as Africa.

There are also some fantastic new views of Saturn’s “sponge moon” Hyperion. The unusual cratering on this small potato-shaped moon makes it look like a giant sponge, but it’s density is also similarly low, about half that of water, indicating the interior of the moon is quite porous. This helps to explain why Hyperion looks the way it does, since impactors would tend to compress the surface, rather than excavate it, and most of the material that is blown off the surface never returns. This is Cassini’s last close flyby of this odd moon, before the scheduled end of the mission in 2017. For its final act, Cassini will repeatedly dive through the space between Saturn and its rings before entering Saturn’s atmosphere.

Cassini’s next flyby encounter will be with the moon Dione on June 16, 2015. The spacecraft’s current position is posted and updated here.

More information on the Cassini mission is available here.
 
Falcon Heavy

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Falcon Heavy is a super-heavy lift space launch system developed and operated by Space Exploration Technologies, SpaceX. Using the company’s Falcon 9 launcher as a basis, Falcon Heavy consists of three F9 cores with a total of 27 Merlin engines, topped by a Falcon 9 upper stage. Operated from Vandenberg Air Force Base and the Kennedy Space Center, Falcon Heavy can be used to access a variety of orbits including Low Earth Orbit, Geostationary Transfer Orbit and interplanetary trajectories. The vehicle includes re-usable technologies and aims to re-use its two side boosters and core stage that make guided boost-back maneuvers and propulsive landings to be refurbished with minimal effort.

Using a standard Falcon 9 and clustering two additional cores to it, Falcon Heavy employs the same overall design principle as the Delta IV Heavy that features three Common Booster Core stages, and theRussian Angara family based on Universal Rocket Modules that can be clustered to cover different payload classes. Falcon Heavy is larger than these two launchers and capable of reaching twice the payload capability of the Delta IV Heavy which currently is the most-powerful space launch system in the world.

SpaceX initiated the development of its heavy launch system in the first half of the 2000s with the goal of creating a launcher that can compete with the world’s heavy lifters such as Delta IV Heavy and Ariane 5 on the commercial launch market, focused on commercial satellites headed to Geostationary Transfer Orbit. Originally, Falcon Heavy was planned to make its first flight two years after Falcon 9, starting out as “Falcon 9 Heavy” since another version based on the now-canceled Falcon 5 was also planned.

Initial performance data for Falcon Heavy was published by SpaceX in 2006 showing a Low Earth Orbit payload capability of 24,750 Kilograms and a launch cost of $78 million. These numbers changed quite often, usually trending up with LEO capabilities rising to 28 metric tons by 2007 and to 32 metric tons by 2010 with a projected launch cost of $95 million. By 2011, Elon Musk announced that development of the launcher was completed, now using the v1.1 version of Falcon 9 as a baseline which further increased the size and performance of the Falcon Heavy.

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Falcon Heavy’s first flight was originally planned out of Vandenberg with an initial target of 2013. After SpaceX took over operation of Launch Complex 39A at the Kennedy Space Center and Falcon Heavy fell behind schedule, its first mission was shifted from the west to the east coast.


As of 2015, Falcon Heavy is shown to have a payload capability of 53,000kg into LEO, 21,200kg to Geostationary Transfer Orbit and 13,200kg that can be inserted into a Trans-Martian Trajectory. Falcon Heavy employs re-usable technologies that are also used on the Falcon 9 such as grid fins, landing legs and re-ignition capability to fly the outer cores and the core stage back to the launch site or a floating platform in the ocean. Depending on the re-usability mode that is selected, Falcon Heavy will be facing payload penalties.

Falcon Heavy surpasses the payload capability of all current launch vehicles and only falls short to the Saturn V and Energia rockets that were capable of carrying more mass into orbit. All requirements for human-rating are being met or exceeded by Falcon Heavy, broadening its potential future applications.


Launch Vehicle Description

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Falcon Heavy stands 68.4 meters tall with a core diameter of 3.66 meters and a total launch mass of 1,462,836 Kilograms consisting of a Falcon 9 core stage with two stretched cores attached to the central stage. The launcher uses a standard Falcon 9 second stage and 5.2-meter diameter payload fairing.

Each of the cores sports nine Merlin 1D engines for a total number of engines on the first stage of 27, only surpassed by the Soviet N1 rocket in terms of the number of engines ignited at liftoff. Propellant crossfeed between the cores is an optional upgrade that will be used for the heaviest payloads (>45mT LEO), otherwise, the central core would throttle down its engines to be able to burn beyond the propulsion phase of the outer cores. The second stage is equipped with a Merlin 1DVac engine optimized for operation in vacuum.

All stages of Falcon Heavy use Rocket Propellant 1 fuel and Liquid Oxygen oxidizer, employing propellant densification to optimize the mass fraction of the vehicle, further pushing payload capabilities and allowing Merlin 1D+ to operate at its full potential.

Core Stage & Merlin 1D Engine



Falcon Heavy uses a central core stage that is nearly identical to that of the Falcon 9 v1.1 (F9R) rocket with the only difference being the addition of interfaces with the outer boosters.

The Core Stage of Falcon Heavy stands about 45.7 meters tall and is 3.66 meters in diameter featuring the standard design with the oxidizer tank located above the fuel tank. Monocoque structure is utilized on the oxidizer tank while the fuel tank features a stringer and ring-frame design that adds strength to the vehicle. The first stage tank walls and domes are made from aluminum lithium alloy and utilize reliable welding techniques to provide maximum strength.

All components of Falcon 9 and Falcon Heavy are designed with structural safety margins 40% above the expected flight loads, higher than the 25% margin that has become the standard in the industry.

The first stage uses Liquid Oxygen oxidizer and Rocket Propellant-1 as fuel which is highly refined Kerosene. The LOX feedline is routed through the center of the fuel tank to supply oxidizer to the engines.

The exact dimensions and mass of the core stage are unknown, but the common belief is that it is capable of carrying about 414,000kg of propellants when prop densification is employed. The stage is about 45.7 meters in length (with interstage), 3.66 meters in diameter and has an empty mass of about 23 to 26 metric tons.

Falcon Heavy sports nine Merlin 1D engines on each of its cores. Compared to its 1C predecessor, Merlin 1D uses improved manufacturing and quality control techniques to enable SpaceX to produce a greater number of engines per year while reducing overall risk. The M1D design is simplified over the M1C by removing no-longer-needed subassemblies. Electro-plating of a nickel-cobalt alloy on the chamber to create the jacket that endures the primary stress of the pressure vessel was replaced by using an explosively formed metal jacket. These changes provide the Merlin 1D with an increased fatigue life and greater thermal margins for the chamber and nozzle which come into play when operating the M1D in an enhanced setting, here referred to as M1D+.

Merlin 1D is an open-cycle gas generator engine. The gas generator operates fuel-rich, burning a small fraction of the LOX and RP-1 flow from the turbopumps to generate a hot high-pressure gas that drives a single turbine with the two turbopumps being driven by a single shaft. High-pressure RP-1 from the fuel turbopump is used in the hydraulic actuators that gimbal the nine main engines for thrust vector control. Generator gas flows through a heat exchanger which heats up Helium gas for tank pressurization in flight before the generator gas is being dumped overboard through an exhaust. The Kerosene flow from the pump is directed to the combustion chamber and nozzle where it passes through heat exchangers as part of the regenerative cooling scheme of the engine. After passing through the heat exchangers, the fuel is pumped into the combustion chamber where it comes into contact with the oxidizer. Merlin 1D operates at a high chamber pressure of 97bar to generate a sea level thrust of 654 Kilonewtons (66,700kg) and a vacuum thrust of 716kN (73,000kg) - giving Falcon Heavy a total liftoff thrust of 17,615kN (1,796,230 Kilogram-force). Vehicle control is provided by gimbaling the nine Merlin engines when the core stage is on its own, the outer boosters can also individually gimbal their engines.

The engine has an increased expansion ratio of 16 while the M1C engine had an expansion ratio of 14.5. Merlin 1D achieves the a thrust to weight ratio of 155 - the highest thrust-to-weight ratio in the liquid-fueled engine world. Merlin 1D uses a pyrophoric mixture of Triethylaluminum-Triethylborane (TEA-TEB) as igniter that is injected into the gas generator and combustion chamber to initiate the combustion process that is sustained as LOX and RP-1 flows into the GG/Chamber once turbopumps spin up, initially using high-pressure helium for spin-up.

Also, the engine has a deep throttling capability which allows Falcon to fly a flexible mission profile. The baselined throttle capability ranges from 70 to 112% of rated performance, however, there are strong indications that M1D can throttle down to 40 or even 30%. To facilitate the propulsive return of the cores, a subset the Merlin 1D engines of the first stage feature onboard re-ignition systems to be fired several times in flight.
All three cores of the Falcon Heavy feature the "Octaweb" engine arrangement. Eight engines are arranged in a circle - clustered around a single Merlin 1D in the center that is installed slightly lower with its nozzle protruding the others. The gas generator exhaust pipes of the individual engines installed on the perimeter of the first stage are arranged toward the inboard direction, their flow passing through the gap between the center and the outer engines, transporting excess heat out of the engine compartment.

The skin of the launcher is the primary load path for the launch vehicle and arranging most of the engines on the perimeter of the skin eliminates a lot of structure that needs to be installed to carry loads from the engines to the skin. The original tic-tac-toe engine pattern required these load-transferring structures, adding to the overall mass of the vehicle. The new engine arrangement also improves thermal properties as it avoids hot spots.

Core Stage

Type Falcon 9 v1.1 Stage 1
Length* 45.7m
Diameter 3.66m
Inert Mass* 25,600kg
Propellant Mass* 414,000kg
Fuel Rocket Propellant 1
Oxidizer Liquid Oxygen
RP-1 Mass* 124,000kg
LOX Mass* 290,100kg
LOX Tank Monocoque
RP-1 Tank Stringer & Ring Frame
Material Aluminum-Lithium
Guidance From 2nd Stage
Tank Pressurization Heated Helium
Propulsion 9 x Merlin 1D+
Engine Arrangement Octaweb
Engine Type Gas Generator, Open-Cycle
Propellant Feed Turbopump
M1D+ Thrust (100%) Sea Level: 654kN - Vac: 716kN
Engine Diameter ~1.0m
Engine Dry Weight 450 to 490kg
Burn Time* 260s
Specific Impulse 282s (SL) 311s (Vac) [for M1D]
Chamber Pressure 108 bar
Expansion Ratio 16
Throttle Capability 70% to 112% (Possibly Deeper)
Restart Capability Yes (Partial)
Ignition TEA-TEB
Attitude Control Gimbaled Engines (pitch, yaw, roll)
Cold Gas Nitrogen RCS
4 Grid Fins (S1 Interstage)
Shutdown Commanded Shutdown
Stage Separation Pneumatically actuated
mechanical collets

Merlin 1D+ & Propellant Densification

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Using improved manufacturing techniques and materials, the Merlin 1D engine was developed with a great margin in operational conditions and a high degree of durability which would enable the engine to operate at higher thrust levels, pressures and temperatures than originally envisioned. In a 2013 press briefing, Elon Musk stated that Merlin 1D could be operated at a sea level thrust of 734 Kilonewtons, representing about a 12% increase in thrust (other unconfirmed numbers that floated around indicated thrust increases up to 20%). Perhaps SpaceX was already looking toward Falcon Heavy when designing the Merlin 1D for operation at this increased thrust setting. Running the engine at a greater propellant mass flow rate will lead to a higher chamber pressure and combustion temperature, increasing the overall stress on the engine. When flying on Falcon Heavy, this M1D+ engine would set up the proper initial thrust to weight ratio and reduce gravity losses in the early ascent phase.

Another technique to be employed by Falcon Heavy is propellant densification for an improved propellant mass fraction and a longer burn time with associated performance increase.

Elon Musk stated that propellant densification capability would be added to all SpaceX launch facilities and it is likely that all Falcon Heavy missions will rely on densification. Densifying propellants is possible through cooling – increasing the mass than can be loaded into the limited tank volume of the launcher. NASA studies have shown that LOX densification can increase the oxidizer mass by 8 to 10% compared to boiling-point LOX at –183°C. Cooling LOX below its boiling point is possible through the use of a Nitrogen subcooler that employs a Liquid Nitrogen bath (either at boiling point or sub-cooled) through which the LOX lines are running to allow an exchange of heat.

LOX temperatures of below –200°C are achievable, however, an economic consideration is necessary when choosing the desired LOX temperature. Operational launchers that employ sub-cooled LOX are Antares (in its original version, using LOX at –196°C) and Soyuz 2-1v (-192°C LOX). Using sub-cooling, the oxidizer mass held in the central core stage's tank of Falcon Heavy could be increased by ~24,000 Kilograms.

Sub-cooling the fuel, Rocket Propellant 1, is also possible, although its high freezing temperature of approximately –37°C and changes in viscosity as a function of temperature represent limitations when sub-cooling the fuel. Through sub-cooling the RP-1 to –25° to –30°C, the first stage could take on ~6 metric tons of additional fuel which will not match up with the increase in LOX. This could be compensated by filling an unused portion in the RP-1 tank (if part of the design) or have the M1D+ engines operate less fuel-rich, leading to increased temperatures and raising demand on the regenerative cooling system of the engine. Equipment for propellant densification including LOX subcoolers has been spotted at the SpaceX launch pad at Vandenberg Air Force Base in 2014 and the addition of this type of equipment at SLC-40 at CCAFS and LC-39 at KSC is likely planned or already in progress.


Just like the Falcon 9, FH provides engine-out capability for a large portion of its first stage flight. All 27 engines are ignited on the ground, about three seconds before launch. All must reach operational conditions and liftoff thrust for the launch release command to be issued.

The engines are monitored constantly in flight and computers can shut down any engine at any time to prevent RUD (rapid unplanned disassembly). Following the unplanned shutdown of an engine, the flight computer would re-plan the ascent trajectory to reach the cutoff target with the remaining engines by extending their burn and potentially cutting a booster return, forgoing re-usability to ensure success of the primary mission.

Flying 9 Merlin engines per core provides engine-out capability and it also allows Merlin 1D to quickly build up flight heritage as each mission provides performance data on 27 engines instead of a single engine that competing launchers are using.The core stage and boosters are equipped with a cold-gas Reaction Control System using Nitrogen for three-axis control during coast phases and during single-engine burns.

The Falcon Heavy cores employ an S-Band communications system to transmit performance telemetry throughout the flight and after stage separation. Each core is equipped with a Flight Termination System consisting of two strings of transmitters, receivers and safe and arm devices. The FTS works with C-Band Communications and can be used to terminate the flight in case of any major anomalies.

The core stage of the Falcon Heavy is connected to the second stage via a carbon fiber aluminum core composite structure acting as interstage adapter, housing the MVac engine of the second stage. Stage separation is accomplished via separation collets and pneumatic pushers in three interfaces connecting the two stages. SpaceX tries to avoid using pyrotechnics for separation events.

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Falcon Heavy Boosters

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Like the Delta IV and Angara, Falcon Heavy aims to reduce complexity in its design by using as much commonality between its core stage and the strap-on boosters as possible. For Delta IV and Angara, the boosters are the same dimension as the core, only using aerodynamic caps and attachment mechanisms that make them different from the core. For Falcon Heavy, some complexity is added by stretching the outer cores as compared to the central core. (No specifics are available on this, however, all animations and scale drawings of Falcon Heavy clearly show the boosters being around six meters taller than the central core (without interstage). Additionally, adding up the masses of the components leaves a much greater mass for the boosters than the estimated mass of the F9 core.)

The boosters of Falcon Heavy share the 3.66-meter diameter of the core stage which allows the same tools and techniques to be used in the manufacturing process, the only difference in the structural design being the stretched tanks. Each of the boosters is approximately 48 meters in length and weighs around 470 metric tons when fully fueled for launch. The boosters each use nine Merlin 1D engines also arranged in an Octaweb pattern and each core is outfitted with independent Guidance, Navigation and Control Systems with communication paths between the computers of the central core and the boosters to allow the main flight computers to issue commands to the boosters and separation systems.

Atop each of the boosters sits a nosecone manufactured from composite materials to keep its weight at a minimum. The four grid fins of the boosters are installed in the uppermost portion of the propellant tank structures, matching up in height with the fins of the central core which reside on the interstage. Each booster has its own nitrogen cold gas reaction control system and is capable of executing an autonomous return to the launch site to be re-used.

The boosters are attached to the central core stage via structural interfaces in the aft section and interfaces that connect the upper portion of the boosters to the interstage area of the Falcon Heavy via thrust struts to transfer loads to the vehicle. Separation of the boosters is accomplished using collets in the structural interfaces, avoiding the use of pyrotechnics since SpaceX prefers to use systems that can be tested and re-used. The reaction control system of the boosters ensures a clean separation from the core stage.

Falcon Heavy Boosters

Type Falcon Heavy Booster
Length* ~47.7m
Diameter 3.66m
Inert Mass* 26,500kg
Propellant Mass* 443,000kg
Fuel Rocket Propellant 1
Oxidizer Liquid Oxygen
LOX Mass* 310,800kg
RP-1 Mass* 132,200kg
LOX Tank Monocoque
RP-1 Tank Stringer & Ring Frame
Material Aluminum-Lithium
Guidance From 2nd Stage
Tank Pressurization Heated Helium
Propulsion 9 x Merlin 1D+
Engine Arrangement Octaweb
Engine Type Gas Generator, Open-Cycle
Propellant Feed Turbopump
M1D+ Thrust (100%) Sea Level: 654kN - Vac: 716kN
Engine Diameter ~1.0m
Engine Dry Weight 450 to 490kg
Burn Time* 190s
Specific Impulse 282s (SL) 311s (Vac) [for M1D]
Chamber Pressure 108 bar
Expansion Ratio 16
Throttle Capability 70% to 112% (Possibly Deeper)
Restart Capability Yes (Partial)
Ignition TEA-TEB
Attitude Control Gimbaled Engines (pitch, yaw, roll)
Cold Gas Nitrogen RCS
4 Grid Fins
Shutdown Commanded Shutdown
Stage Separation Thrust Struts, RCS

Throttle-Down vs. Crossfeed

When Falcon Heavy was initially announced, one of its biggest innovations was to be a propellant crossfeed capability between the boosters and the central core. The design called for propellant lines being routed through the interfaces of the cores, delivering propellants from the oxidizer and fuel manifolds of the boosters to a number of engines on the core, allowing these engines to consume propellant from the booster tanks and leaving the tanks of the central core nearly full until the point of separation when interfaces would be isolated and supply switch to the core stage’s own tanks.

This feature was put on the backburner and SpaceX decided to introduce the Falcon Heavy without operational crossfeed system using a partial thrust mode on the core stage to allow it to save propellants that can be consumed beyond the burn of the side boosters. This procedure is also used by Delta IV Heavy and the Angara family, eliminating the additional mass of a crossfeed system that would only be required on flights with extremely heavy payloads. In a nominal flight scenario, Falcon Heavy would take off with all of its Merlin 1D engines at full throttle, likely to be the 112% setting as a standard. After the initial climb, the central core would throttle its engines down to a minimal thrust in order to save propellants while the boosters continue to fire at full thrust.

The two boosters would burn for roughly 180-200 seconds before separating from the core to begin their journey back to the launch site. Continuing powered ascent, the core would throttle up its engines and burn for just over a minute to continue boosting the velocity of the stack, achieving a much higher speed than any previous Falcon cores which makes its recovery more difficult given its much greater energy at separation.

SpaceX is still pressing ahead with the development of the Crossfeed Capability to be used on particular challenging missions with LEO payloads of over 45 metric tons or equivalent payloads to different orbits.

Re-Usable Falcon Heavy

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Re-usability of space launch vehicles is one of the biggest goals of SpaceX and Falcon Heavy aims to become the first partially re-usable super-heavy lift launcher. Like Falcon 9, FH will return its three cores to the ground through a series of propulsive maneuvers, a guided flight through the atmosphere and a soft landing on four deployable landing legs – either on land using a flat landing pad to be built near the launch site or on the Autonomous Spaceport Drone Ship inaugurated during the initial tests of returning Falcon 9 boosters to an on-target landing.

The overall goal is to get the rocket stages back to the launch site to avoid the cost of having them returned by ship or other means of transportation. This boost-back to the launch site is feasible for the two outer cores that separate from the launch vehicle at a much lower energy than the central core. Continuing powered flight, the core stage reaches a high energy that would cause a large payload penalty due to the additional fuel required for the boost back to the launch site.

Therefore, SpaceX will keep operating the drone ships (one for the east coast, one for Vandenberg launches and potentially another one for the Brownsville launch site) for the return of the core stages. SpaceX was also looking into the possibility of refueling the cores on the landing platform and having them fly back to land under their own engine power.

In an operational scenario, Falcon Heavy would blast off and burn its two outer boosters for close to three minutes before the two boosters separate and begin their journey back to the launch site. Immediately, the two boosters would rotate to an engines-first position and make their way to apogee for the boost back burn that would use a subset of the Merlin 1D engines.

This boost back would reverse the downrange velocity and allow the boosters to begin traveling back to the launch site. Passing through 70 Kilometers in altitude, the boosters would ignite three of their engines for the re-entry burn that serves two purposes – starting to slow the booster down and providing protection to the engine compartment from the aerodynamic re-entry environment. Falcon 9 re-entry burns had a typical duration of 19 seconds.

Beginning their descent through the atmosphere, the boosters would deploy four grid fins for precise steering.

The four grid fins are launched in a position stowed against the uppermost section of the booster near the nose cone before being deployed when Falcon re-enters the atmosphere. The four fins can be individually controlled in a two-degree of freedom type design, rotating and tilting at the same time, allowing for complex guidance and control during atmospheric flight.

The fins are an essential part of Falcon’s return sequence to provide control in atmospheric flight without active propulsion. Grid-fins have been widely used as a stabilizer on missiles & bombs and are shaped like miniature wings consisting of a lattice structure. The Russian Soyuz employs grid-fins in its launch abort system which would deploy when the launch escape rockets start firing in an abort scenario to stabilize the vehicle, but the fins used by SpaceX take it one step further as they can be moved independently to actively control the vehicle's flight and not only act as a stabilizer.

Grid-fins perform well in all velocity ranges including supersonic and subsonic speeds with the exception of the trans-sonic regime due to the shock wave enveloping the grid. These properties make them ideally suitable for the Falcon booster stages that start out at supersonic speeds and return to subsonic velocity as they travel through the atmosphere, en-route to the landing site. The four fins are rotated and tilted independently by an open hydraulic system that uses pressurized hydraulic fluid supplied from a pressurized tank that is dumped overboard after flowing through the hydraulic actuators of the fin system. The design was also driven by overall mass considerations.

The addition of the grid fins was expected to improve the accuracy of Falcon’s landing by three orders of magnitude – previous landing attempts in the ocean had a ten-Kilometer targeting accuracy while the return to a platform or a pad on land requires the stage to land within a few meters of its bulls-eye target.

Heading back in, the boosters would make final corrections to their flight path, modifying their pitch trim to precisely target their landing site. Around 28 seconds prior to touchdown, the center engine of the booster is re-ignited for the final landing burn. With a limited throttle range, the Center engine will generate a thrust that is greater than the mass of the stage. Landing at a thrust to weight ratio greater than one requires the stages to calculate their propulsive landing maneuver in a way that that reaches a minimum velocity when coming into contact with the ground. Falcon’s boosters are targeting to land at a velocity of less than 6 meters per second.

Ten seconds prior to touchdown, the four landing legs of the booster would deploy. The overall design driver for the landing legs was mass since adding significant weight to the first stage would have resulted in a significant payload penalty. Safety was also a major concern – the leg design had to be such that no premature deployment during powered ascent was possible which would result in a certain loss of the entire vehicle and payload.

Made of aluminum honeycomb and carbon-composite materials, the four legs have a total mass of around 2,100 Kilograms consisting of a single-load bearing strut and aerodynamic fairing assembly. The central struts of the legs interface with the load-carrying structure of the first stage while the fairings have two structural interfaces at the base of the engine compartment heat shield and one interface on the lower portion of the leg
During flight, the legs are stowed against the rocket body, covered by the fairings that ensure no additional aerodynamic disturbance is introduced by the legs. Deployment is accomplished by a pneumatic system using high-pressure helium. When deployed, the legs have a span of about 18 meters, capable of supporting the forces of landing and the mass of the nearly empty booster.

SpaceX has secured properties at Cape Canaveral and Vandenberg Air Force Base to be used as booster landing facilities. At Cape Canaveral, SpaceX signed a five-year lease of Launch Complex 13 in February 2015. An animation of the Falcon Heavy flight profile shows a conceptualized representation of LC-13. Five individual flat landing pads are seen in the animation with four smaller auxiliary pads and one larger central pad. The two boosters use two of the smaller pads, landing within seconds of each other after making their propulsive return from the edge of space.

LC-13 at CCAFS has been in operation from 1958 to 1978 supporting the Atlas launcher family with notable LC-13 launches including Lunar Orbiter 1 and a number of Atlas Agena launch vehicles. The launch pad was not in use for nearly three decades and had its mobile service tower demolished in 2005 followed by the demolition of the blockhouse in 2012. LC-13 will be used to return Falcon stages launching from SLC-40 and LC-39.

At Vandenberg Air Force Base, SpaceX has procured Space Launch Complex 4W for booster landings with SLC-4E serving as Falcon 9 and Falcon Heavy launch pad.

SLC-4W was active for over four decades starting in 1963, supporting Atlas-Agena missions before being converted for the Titan II launch vehicles. In total, SLC-4W saw over 90 launches before becoming inactive after the last Titan 23G launch in 2003. In 2014, the complex was handed to SpaceX and the demolition of existing structures including the Mobile Service Tower started in September 2014. The finished landing facility will likely look very similar to that at Cape Canaveral.

Due to the central core stage of Falcon Heavy continuing onwards after booster separation, it faces a much higher speed when separating from the second stage. A return to the launch site would require a considerable amount of propellant leading to a large payload penalty. Therefore, SpaceX will keep using the Autonomous Spaceport Drone Ships that will be stationed downrange from the launch site to welcome the core stages. The downrange distance of the drone ship will depend on the surplus of propellant that is available for an active boost back.
Known as the Autonomous Spaceport Drone Ship, the floating landing platform was built at a Louisiana shipyard and measures 91 meters by 52 meters with a prominent Space“X marks the Spot” logo in the center. The ship sports four diesel-powered azimuth thrusters – similar to those on oil rigs - provided by Thrustmaster, a marine equipment manufacturer that also provided power modules and controls to outfit the ship with a Portable Dynamic Positioning System. Processing GPS data, the Autonomous Spaceport Drone Ship will be able to keep its assigned position with an impressive accuracy of three meters.

A high accuracy is required since Falcon will have to land on the platform with all four of its legs that span approximately 18 meters, leaving just over 30 meters for GPS errors between the two craft and position errors of the drone ship, sea swell as well as errors by Falcon, making its fast-paced hoverslam landing under the power of one of its nine Merlin 1D engines with a thrust to weight ratio greater than one.

The ASDS is outfitted with a water deluge system that dumps water onto the deck to protect it from the heat of the engine of the arriving booster. Numerous attachment fixtures are part of the deck structure that would allow the securing of the first stage after landing on the platform for the return to port and refurbishment.

Flying as fully expendable launch vehicle, Falcon Heavy could deliver 21,200 Kilograms to a standard Geostationary Transfer Orbit. With full reusability on all three cores, the launcher will only be able to put seven metric tons into GTO which is still within the mass range of the heaviest commercial communications satellites. Only returning the boosters and flying the central core as expendable booster will increase GTO capability well over ten metric tons.

Second Stage

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Falcon Heavy uses a standard Falcon 9 v1.1 second stage, potentially employing propellant densification to optimize launch vehicle performance.

The second stage of the Falcon 9 is based on the design of the v1.0 second stage which is essentially a smaller version of the first stage. SpaceX has always followed a policy of choosing simple solutions to reduce cost and risk in order to manufacture a robust launch system. Using the same materials, tools and manufacturing techniques for the two stages is a perfect example of this approach.

As with the first stage, the exact dimensions of the second stage have not yet been disclosed by SpaceX. It is estimated that the second stage is 15 meters long with an inert mass of around four metric tons and a fuel load of 97,000 Kilograms. The diameter is identical to the core stage.

Comparing it with the v1.0, the second stage of the v1.1 features stretched propellant tanks that are also built using Aluminum-Lithium Monocoque structure for both tanks. The second stage also uses Rocket Propellant 1 as fuel and Liquid Oxygen as oxidizer.

One Merlin 1D Vac engine is powering the second stage. This engine differs from the first stage engines as it is optimized for operation in vacuum featuring an extended nozzle with a high expansion ratio. M1D Vac is also a turbopump-fed gas generator engine, in its enhanced version, it operates at a chamber pressure of 108 bar.

Using an extended nozzle creates a high expansion ratio of greater than 117:1. M1D Vac has a high specific impulse of over 340s that could be as high as 347s. It generates a total vacuum thrust of 801 Kilonewtons (81,700 Kilograms) when flying as standard M1DVac, the M1D+ Vacuum version could achieve a thrust of 897kN. The engine can support multiple ignitions to be able to fly a flexible mission profile in order to reach a variety of orbits and trajectories. The second stage TEA-TEB ignition system is fully redundant.

Second Stage Burn time is variable with nominal firings of ~372 seconds.

The second stage is equipped with a Reaction Control System for three axis-control during coast phases and roll control during burns. The Falcon 9 v1.1 uses a cold-gas attitude control system employing a number of Nitrogen thrusters for three axis control during extended coast phases.

The second stage of the Falcon rocket facilitates the avionics and flight computers that control all aspects of the flight. The avionics of the Falcon feature a number of changes and upgrades from the v1.0 to the v1.1 and Heavy version. All avionics and controllers are manufactured in-house by SpaceX. The system is fully redundant, constantly checking itself to verify that all GNC components are functioning properly. SpaceX uses commercial off-the-shelf parts that are radiation tolerant instead of radiation hardened (cost reduction). The flight computers run on Linux with software written in C++.

Avionics are triple redundant and the rocket’s inertial navigation system uses GPS overlay for additional orbital insertion accuracy.

In addition to the main avionics units of the launch vehicle, each of the Merlin Engines is equipped with three processing units in a single engine controller. The engine controller monitors all parameters of the engine and interfaces with the main avionics units. Each of the three processing units are constantly checking on the others to provide fault-tolerance.

Second Stage

Type Falcon 9 v1.1 Stage 2
Length* 15.2m
Diameter 3.66m
Inert Mass* 4,000kg
Propellant Mass* 97,000kg
Fuel Rocket Propellant 1
Oxidizer Liquid Oxygen
LOX Mass* 68,800kg
RP-1 Mass* 28,200kg
LOX Tank Monocoque
RP-1 Tank Monocoque
Material Aluminum-Lithium
Guidance Inertial
Tank Pressurization Heated Helium
Propulsion 1 x Merlin 1D Vac +
Engine Type Gas Generator
Propellant Feed Turbopump
Thrust 897kN (M1D+)
Engine Dry Weight 450 to 490kg
Burn Time* 372s
Specific Impulse >340s (Est: ~345s)
Chamber Pressure 108 bar
Expansion Ratio >117
Throttle Capability Yes
Restart Capability Yes
Ignition TEA-TEB, Redundant
Pitch, Yaw Control Gimbaled Engine
Roll Control Reaction Control System
Shutdown Commanded Shutdown
Reaction Control S. Cold-Gas Nitrogen Thrusters

Payload Fairing

The Payload Fairing is positioned on top of the stacked vehicle and its integrated spacecraft. It protects the vehicle against aerodynamic, thermal and acoustic environments that the launcher experiences during atmospheric flight. When the launcher has left the atmosphere, the fairing is jettisoned. Separating the fairing as early as possible increases ascent performance.

Falcon 9's standard Fairing is 13.1 meters in length and 5.2 meters in diameter. The fairing consists of an aluminum honeycomb core with carbon-fiber face sheets fabricated in two half-shells. Separation is accomplished via a pneumatic system along the vertical seam that pushes the two halves apart.
Up to three spacecraft access doors or Radio Frequency Windows can be supported by the fairing. A small 3.6-meter fairing is also being developed.


Payload Adapters

Payload Adapters interface with the vehicle and the payload and are the only attachment point of the payload on the Launcher. They house equipment that is needed for Spacecraft Separation and ensure that the payload is secured during powered flight. Electrical and Communication connections are also part of the Adapter and route spacecraft Telemetry to the Flight Computers for downlink. A variety of different adapters is available to suite different spacecraft needs and requirements.

EO Injection Accuracy (v1.0)

Perigee +/- 10km
Apogee +/- 10km
Inclination +/- 0.1 deg
Right Ascension of
Ascending Node +/- 0.15 deg

GTO Injection Accuracy (v1.0)

Perigee +/- 7.4km
Apogee +/- 130km
Inclination +/- 0.1 deg
Right Ascension of
Ascending Node +/- 0.75 deg
Arg of Perigee +/- 0.3 deg

Payload Fairing

Payload Fairing Composite Fairing
Diameter 5.2m
Length 13.1m
Weight ~1,750kg
 
Minotaur V


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The Minotaur V is an expendable launch system developed and operated by Orbital Sciences Corporation and the US Air Force. The launcher is based on the all-solid Minotaur IV launch vehicle which itself is based on the Peacekeeper Intercontinental Ballistic Missile.

Peacekeeper was a land-based ICBM that began development in 1972. At the time, silo-based Minuteman ICBMs were being deployed by the US. The development of the R-36M missile by the Soviets gave the Soviet Union the theoretical the ability to destroy the US Minuteman ICBM facilities before retaliation would have been possible. This prompted the development of the Peacekeeper launcher that could also deliver warheads to orbit via a Post Boost Vehicle and Deployment Module to independently target the individual warheads. The first test launch of the Peacekeeper took place in 1983 from Vandenberg Air Force Base in California and the system was deployed in 1986. In 2003, the retirement process of the Peacekeeper was started and by September 2005, the last Peacekeeper was removed from alert status.

Peacekeeper warheads are being deployed on Minuteman III missiles that is the only land-based ICBM currently in use by the United States. The Peacekeeper rockets themselves are being converted to orbital launch vehicles by Orbital Sciences. The rockets are converted by fitting them with a fourth stage (& optional 5th & 6th stages) and by installing Orbital’s enhanced avionics systems and advanced composite structures to facilitate its payloads. Combining the robust heritage components flown on the Peacekeeper and advanced avionics and support systems create a low-cost launch vehicle for use to support government-financed launches.

To make the launch systems more flexible, Orbital Sciences developed different versions of the basic four-stage Minotaur IV. In its Minotaur IV+ configuration, the vehicle uses a the more powerful Star-48V upper stage instead of the Orion-38. Flying entirely without a fourth stage, Minotaur IV Lite can be used for sub-orbital flights.

In the Minotaur V configuration, a Star 37 rocket stage is added to the IV+ four-stage stack for launches to trans-lunar trajectories and Geosynchronous Transfer Orbit. A six-stage Minotaur VI version has also been conceptualized.

Minuteman launch vehicles are operated from Space Launch Complex 8 at Vandenberg Air Force Base, Launch Pad 1 at Kodiak Launch Complex (Alaska) and Pad 0B at the Mid-Atlantic Regional Spaceport (MARS), Virginia.

Minotaur IV has launched five times starting in April 2010. The Minotaur V version has not flown and its first launch is planned to be to deliver NASA’s LADEE spacecraft to a trans-lunar trajectory.

Launch Vehicle Description

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The Minotaur V launch vehicle stands nearly 24.5 meters tall with a diameter of 2.34 meters and a liftoff mass of about 89,000 Kilograms. It uses the three stages of the Peacekeeper rocket designated SR-118, SR-119 and SR-120. These stages are provided by the US government and are used without any major modifications. Minotaur V uses two solid-fueled upper stages manufactured by Alliant Techsystems making it an all-solid launch vehicle. The fourth stage is an ATK Star 48 BV rocket motor and the fifth stage is the smaller Star-37 that can be flown in two different configurations.

Minotaur V can deliver small payloads to a variety of trajectories including Medium Earth Transfer Orbit, Geosynchronous Transfer Orbit and Trans-Lunar Trajectories.

Minotaur V Specifications

Type Minotaur V
Manufacturer Orbital Sciences
Operator OSC, USAF
Launch Site Vandenberg, Kodiak, MARS
Height ~24.5m
Diameter 2.34m
Launch Mass ~89,000kg
Stages 5
Stage 1 SR-118
Stage 2 SR-119
Stage 3 SR-120
Stage 4 Star 48 BV
Stage 5 Star 37 (FM or FMV)
Mass to GTO 532kg
Mass to MTO 650kg (CCAFS), 603kg (WFF)
Mass to TLI 342kg


First Stage

The first stage of the Minotaur V rocket is the first stage of the Peacekeeper that is flown without major modifications. SR-118 was manufactured by Thiokol and is also known as TU-903 and uses HTPB (Hydroxyl-terminated polybutadiene) based propellants. The first stage is loaded with 45,400kg of propellant that is consumed during the 56.5-second burn of the stage to provide 2,224 Kilonewtons of thrust (226,780 Kilograms).

Control during first stage flight is provided by a hydraulic Thrust Vector Control System steered with actuator commands provided by the Booster Control Module that links the flight computer to the TVC system.

First Stage

Type SR-118 (TU-903)
Diameter 2.34m
Length 8.4m
Propellant Solid - HTPB
Launch Mass 49,000kg
Empty Mass 3,600kg
Propellant Mass 45,400kg
Guidance via Booster Control Module
Propulsion TU-903
Thrust 2,224kN
Burn Time 56.5sec
Specific Impulse 229sec (SL), 284sec (Vac)
Control Hydraulic Thrust Vector Control


Second Stage

The second Stage of the Minotaur launcher was manufactured by Aerojet and also uses HTPB-based propellant. It is 2.34 by 7.9 meters in size with a launch mass of 27,700 Kilograms. It closely resembles the design of the first stage and also uses a hydraulic Thrust Vector Control System that provides attitude control during the 61-second burn of the second stage. It provides a thrust of 1,223 Kilonewtons (124,710 Kilograms).

Second Stage

Type SR-119
Diameter 2.34m
Length 7.9m
Propellant Solid - HTPB
Launch Mass 27,700kg
Empty Mass 3,200kg
Propellant Mass 24,500kg
Engine With extendable Exit Cone
Guidance via Booster Control Module
Thrust 1,223kN
Burn Time 61sec
Specific Impulse 308sec (Vac)
Control Hydraulic Thrust Vector Control


Third Stage

The SR-120 served as third stage of the Peacekeeper and is also used as the Minotaur third stage. SR-120 was manufactured by Hercules and uses NEPE propellant containing HMX with greater energy than ammonium perchlorate that is used in most composite HTPB propellants. Propellants containing HMX are not used on commercial launchers because of its explosive hazards, but as a converted ballistic missile, Minotaur uses the unmodified SR-120 with NEPE propellant.

SR-120 is 2.34 meters in diameter, 2.44 meters long and has a total mass of 7,700 Kilograms. It burns for 72 seconds and provides 289kN of thrust (29,470 Kilograms). It also uses a hydraulic thrust vector control system to provide attitude control during its burn.

After the third stage burn, the Minotaur usually performs a coast phase to reach higher altitudes so that the following upper stage burns can serve as circularization maneuvers and raise the perigee of the sub-orbital trajectory to achieve orbit.

Third Stage


Type SR-120
Diameter 2.34m
Length 2.44m
Propellant Solid - NEPE
Launch Mass 7,720kg
Empty Mass 650kg
Propellant Mass 7,080kg
Engine With extendable Exit Cone
Guidance via Booster Control Module
Thrust 289kN
Burn Time 72sec
Specific Impulse 300sec (Vac)
Control Hydraulic Thrust Vector Control

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Star 48BV

The fourth Stage of the Minotaur V launcher is the Star-48 Solid Rocket Motor built by Alliant Techsystems. The ATK Star 48BV is a solid-propellant upper stage that uses the flight proven Star 48B and adds Thrust Vector Capability (V). Star 48 was introduced in 1982 and has been used on a variety of spacecraft. Star 48B was spin stabilized and had smaller a performance than the TVC capable version.

The upper stage features a 1.24-meter diameter titanium casing holding a total of 2,010 Kilograms of solid propellant. It is 2.08 meters in length and has a launch mass of 2,165 Kilograms. It operates at an average thrust of 68.6 Kilonewtons (6,995kg) with peak thrust reaching 77.8kN (7,930kg). Star-48BV features the longer of two available nozzles for the conventional Star 48. The upper stage features an electromechanically actuated flexseal nozzle Thrust Vector Control System with a maximum nozzle gimbal of four degrees. Star-48 burns for 84 seconds.

Star 48BV is the final stage of the Minotaur IV+ launcher and is capable of relatively precise insertions. On the Minotaur V, an additional fifth stage is installed to improve performance for highly elliptical and trans-lunar trajectories.

Star 48BV

Type Star 48BV
Launch Mass 2,164.5kg
Diameter 1.24m
Length 2.08m
Propellant TP-H-3340
Propellant Mass 2,010.0kg
Casing Mass 58.3kg
Case Material Titanium
Nozzle Mass 52.6kg
Avg Thrust 68.6kN
Max Thrust 77.8kN
Isp 288s
Throat Diameter 0.1011m
Nozzle Diameter 0.7475m
Chamber Pressure 39.9bar (Avg) - 42.6bar (Max)
Expansion Ratio 54.8
Burn Time 84.1s
Ignition Delay 0.100s
Attitude Control TVC +/-4°
Roll ACS

Star 37

The Star 37 Solid Rocket Motor is also built by Alliant Techsystems and is the predecessor to Star 48. Two versions of the Star 37 can be used atop the Minotaur V. The Star 37FM version is spin-stabilized while the FMV version is equipped with a three-axis attitude control system. The extra-weight of the control equipment on the FMV version reduces payload capability.

Star 37 FMV weighs 1,170 Kilograms including 1,064kg of TP-H-3340 propellant, but the propellant load can be slightly adjusted based on payload and insertion requirements. It is 0.93 meters in diameter and 1.912 meters long featuring a Nozzle Assembly that uses a 3D carbon-carbon throat and a carbon-phenolic exit cone. Star 37 provides 48.8 Kilonewtons (4,975kg) of average thrust and 55.6kN of peak thrust (5,670kg). Like Star 48, the flexseal nozzle can be gimbaled by up to 4 degrees by the electromechanical Thrust Vector Control System.

A Cold Gas Reaction Control System is used to provide roll control during the upper stage burns and three-axis control during coast phases.

Star 37FM

Type Star 37FM
Launch Mass 1,148kg
Diameter 0.93m
Length 1.69m
Propellant TP-H-3340
Propellant Mass 1,065.9kg
Casing Mass 32.25kg
Case Material Titanium
Nozzle Mass 34.02kg
Avg Thrust 47.3kN
Max Thrust 54.8kN
Isp 290s
Throat Diameter 0.0894m
Nozzle Diameter 0.6215m
Chamber Pressure 37.2bar (Avg) - 44.3bar (Max)
Expansion Ratio 48.0
Burn Time 62.7s
Ignition Delay 0.130s
Stabilization Spin Stabilized - 60RPM

Star 37FMV

Type Star 37FMV
Launch Mass 1,170kg
Diameter 0.93m
Length 1.92m
Propellant TP-H-3340
Propellant Mass 1,063.8kg
Casing Mass 32.25kg
Case Material Titanium
Nozzle Mass 44.91kg
Avg Thrust 48.84kN
Max Thrust 55.6kN
Isp 294s
Throat Diameter 0.0894m
Nozzle Diameter 0.7483m
Chamber Pressure 37.2bar (Avg) - 44.3bar (Max)
Expansion Ratio 70.0
Burn Time 62.7s
Ignition Delay 0.130s
Attitude Control Nozzle TCV - +/-4°
Roll ACS

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Avionics & Guidance System

Minotaur implements a Common Avionics Assembly that is used across the Minotaur family. The CAA is a ring structure that is mounted on the upper stage of the vehicle offering space for the various avionics boxes that comprise the assembly.

The Central Flight Computer of the Minotaur is based on Orbital’ s Modular Avionics Control Hardware (MACH) that provides power transfer, data acquisition, booster interfaces, and ordnance initiation. Up to 10 MACH devices can be combined to satisfy mission requirements. Minotaur’ flight computer uses a 32-bit multiprocessor architecture and a RS-422 serial bus for data connections to avionics and payload systems.

Additionally, the avionics assembly includes the Booster MACH, the Booster Control Module that provides actuator commands to the Thrust Vector Control Systems of the lower stages, the S- and C-Band Communications System that is used for telemetry downlink, Flight Termination System receivers and equipment, a GPS beacon and a vehicle encoder.

Also mounted on the avionics ring is the Attitude Control System of the launcher which is a cold gas system using pressurized Nitrogen. The attitude control system is used for roll control during the 4th and 5th stage burn as well as three-axis control during coast phases and the Contamination and Collision avoidance maneuver.

The Common Avionics Assembly gathers navigation data using an inertial platform that feeds the digital autopilot of the vehicle. The three-axis autopilot is programmed to fly a pre-programmed attitude profile during Stage 1, 2 and 3 flight and gather navigation data which is then used to optimize the trajectory during the Stage 4 & 5 burns.

The two upper stages use a pre-defined set of parameters for their target trajectory which they use to modify their flight profile based on actual achieved trajectory by the lower stages.

The Star 48 and 37 stages uses energy management to achieve the insertion trajectory. After the final boost phase, the three-axis cold-gas attitude control system is used to orient the vehicle for spacecraft separation, contamination and collision avoidance and downrange downlink maneuvers.

Payload Adapters

Minotaur can support a number of Payload Adapter Modules including off-the-shelf adapters and custom built devices. Payload Adapters interface with the launch vehicle and the payload and are the only attachment point of the payload on the Launcher. They provide equipment needed for spacecraft separation and connections for communications between the Upper Stage and the Payload.

A typical PAM consists of a Payload Adapter Fitting that is connected to the upper stage, a payload cone and a separation system. Minotaur can facilitate Orbital-built as well as Planetary Systems and RUAG Space payload attach systems.

Payload Fairing

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Minotaur V uses the standard payload fairing that is also used on the Minotaur IV/IV+ launcher. It is about 6.4 meters long and 2.34 meters in diameter weighing approximately 450 Kilograms. It consists of two composite shell halves, a low-shock frangible rail and ring separation system, and an actuator/hinge fairing jettison system. The fairing structure is a aluminum honeycomb core covered by layers of graphic epoxy composite. The fairing is outfitted with acoustic blankets, a ventilation system and RF windows if required. The fairing also provides access doors to the payloads.

The two fairing halves are joined by a frangible rail joint and the PLF is connected to the second stage using a ring-shaped frangible joint. A cold gas initiation system is used to disconnect the ring and rail so that the two halves of the fairing can rotate outboard on two hinges installed on the vehicle in order to ensure the appropriate clearances during the separation event.

The payload envelope for the Minotaur V is defined by the Star 37 upper stage because it and its support structure has to fit under the fairing together with the spacecraft.

To accommodate larger payloads, Minotaur can be outfitted with a 2.79-meter diameter fairing that features a similar design, but comes at the cost of launch vehicle performance.

Payload Fairing

Type Minotaur IV Fairing
Diameter 2.34m
Length 6.4m (Standard)
Mass 450kg
Separation Ordnance, Frangible Joints,Pistons
Construction Graphite/Epoxy Face Sheets
Aluminum Honeycomb Core
Notes Protects Payload & 5th Stage
 
NASA Awards SpaceX $30 Million for Successful Dragon Pad Abort Test Milestone

NASA Awards SpaceX $30 Million for Successful Dragon Pad Abort Test Milestone Under CCiCap « AmericaSpace

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A mockup SpaceX Crew Dragon takes flight for the company Pad Abort Test at Cape Canaveral Air Force Station on May 6, 2015. Photo Credit: Alan Walters / AmericaSpace

NASA has officially declared SpaceX’s recent Crew Dragon Pad Abort Test (PAT) a success, awarding the Hawthorne, CA-based company $30 million for completion of that very important development milestone under their Commercial Crew integrated Capability (CCiCap) agreement with NASA’s Commercial Crew Program.

The flight test, which took place on May 6, marked a big step forward as SpaceX aims to deliver U.S. astronauts to and from the International Space Station (ISS), aboard a U.S.-manufactured spacecraft, and from U.S. soil, for the first time since the nation’s space shuttle fleet retired from service in 2011.

“This test was highly visible and provided volumes of important information, which serves as tangible proof that our team is making significant progress toward launching crews on American rockets from America soon,” said Jon Cowart, partner manager for NASA’s Commercial Crew Program. “The reams of data collected provide designers with a real benchmark of how accurate their analyses and models are at predicting reality. As great as our modern computational methods are, they still can’t beat a flight test, like this, for finding out what is going on with the hardware.”

Launching off a specially made truss to simulate the spacecraft atop a Falcon-9 rocket from Space Launch Complex-40, the 21,000 pound prototype capsule took flight quickly under 120,000 pounds of axial thrust from its eight SuperDraco engines, which are intended to carry astronauts to safety in the event of an emergency on the pad or during ascent (16,000 pounds of thrust each, compared to 100 pounds of thrust each with the original Draco thrusters on Dragon 1).

The eight SuperDraco engines, which are built directly into Crew Dragon’s walls, are the first fully 3-D printed engines intended for space to ever be developed.

After ascending 3,500 feet in six seconds the PAT Dragon jettisoned its trunk and deployed a pair of drogue chutes, followed by a trio of main parachutes and splashdown less than a mile offshore of the launch site minutes later.

The vehicle was outfitted with hundreds of instruments and sensors for data collection, and even had an instrumented mannequin as the sole passenger, providing SpaceX with important data and other information regarding the stresses put on the mannequin—information that will be critical in ensuring development of an abort system that prevents serious injury to crews.

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SpaceX’s Crew Dragon prototype parachuting back to Earth after a successful Pad Abort Test at Cape Canaveral AFS on May 6, 2015. Photo Credit: John Studwell / AmericaSpace

Dragon’s PAT should provide SpaceX significant data in the areas of Sequencing, Closed-Loop Control, Trajectory, and External and Internal Environments. The PAT demonstrated the proper sequencing of the pad-abort timeline as well, serving to validate the execution of multiple critical commands in a very short period. Trajectory data for both maximum altitude and downrange distance from the pad was gather as well, including data on various internal and external factors to Crew Dragon to help ensure safe conditions for crew transport.

“This is the first major flight test for a vehicle that will bring astronauts to space for the entire Commercial Crew Program,” said Gwynne Shotwell, president of SpaceX. “The successful test validated key predictions as it relates to the transport of astronauts to the space station. With NASA’s support, SpaceX continues to make excellent and rapid progress in making the Crew Dragon spacecraft the safest and most reliable vehicle ever flown.”

The approval of the PAT milestone payment follows NASA’s authorization for Boeing to begin work toward its first post-certification mission with the CST-100 crew capsule, which also received a multi-billion dollar NASA contract for crew transport to and from the ISS. The company recently received the first of up to six orders to execute a crew-rotation mission to the ISS, which NASA stressed does not necessarily imply that a Boeing CST-100 capsule will fly ahead of a SpaceX Crew Dragon.

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Crew Dragon recovered just offshore of its launch site. Photo Credit: Mike Killian / AmericaSpace

SpaceX will conduct one more abort test, an In-Flight Abort atop a Falcon-9 rocket launch, using the same Crew Dragon prototype capsule, later this summer.

Both SpaceX and Boeing are expected to begin carrying out the first operational crewed flights for NASA in 2017, but that is dependent on NASA funding, which is dependent on the federal government. The debate is still ongoing in Congress, but it appears that NASA’s Commercial Crew Program will receive several hundred million dollars less than what the space agency and the White House requested for FY2016, which will likely delay America’s return to human spaceflight from U.S. soil once again.

NASA Administrator Charles Bolden had this to say about it:

“I am deeply disappointed that the Senate Appropriations subcommittee does not fully support NASA’s plan to once again launch American astronauts from U.S. soil as soon as possible, and instead favors continuing to write checks to Russia. By gutting this program and turning our backs on U.S. industry, NASA will be forced to continue to rely on Russia to get its astronauts to space – and continue to invest hundreds of millions of dollars into the Russian economy rather than our own.”

 
@fantastic thread, although wouldn't it be better if I moved it to the US section?
 
although wouldn't it be better if I moved it to the US section?

I stuck it here to encourage casual the use and contribution by members who don't frequent the America's section, which based on the participation in that section seems to be a lot of people, but I'll leave the decision in your hands.
 
I stuck it here to encourage casual the use and contribution by members who don't frequent the America's section, which based on the participation in that section seems to be a lot of people, but I'll leave the decision in your hands.

I'll keep it here my friend.
 
Antares

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Antares is an expendable launch system being developed and operated by Orbital Sciences Corporation. It is a two stage launch vehicle with an optional third stage. The launcher can reach a variety of orbits including Low Earth Orbit, Sun Synchronous Orbit, Geosynchronous Transfer Orbit and Interplanetary Trajectories.

Antares is currently being operated from the Mid-Atlantic Regional Spaceport at NASA's Wallops Flight Facility, however, the vehicle is also compatible with the Western Range at Vandenberg Air Force Base, the Eastern Range at Cape Canaveral Air Force Station and the Kodiak Launch Complex, Alaska.

Orbital Sciences developed Antares under a Commercial Orbital Transportation Services (COTS) contract that the company was awarded from NASA to demonstrate cargo deliveries to the International Space Station. Orbital uses the Antares launcher to boost its Cygnus Capsule to orbit for flights to ISS. Once completing COTS, Orbital enters the Commercial Resupply Services Program. NASA has booked eight ISS resupply flights of Cygnus under Commercial Resupply Services for a total contract volume of $1.9 billion. Antares is also on the market for small and medium missions.

The launch vehicle was originally known as Taurus II, but was renamed in late 2011. Antares made its first flight in April 2013.
The Antares launch vehicle stands 40.5 meters tall, has a main diameter of 3.9 meters and a liftoff mass of approximately 282,000 Kilograms. It uses a liquid fueled first stage that is powered by two powerful engines.

The second stage of the vehicle is a solid-fueled rocket motor built by Alliant Techsystems. As a second stage, the ATK Castor 30A, 30B and XL can be used. For Cygnus missions to ISS, Antares will fly with two stages only, but other payloads may require a third stage.


Two third stages are available for Antares, the Orbital-built Bi-Propellant Upper Stage that allows Antares to perform precise injections into a variety of orbits and the ATK Star 48 Solid Upper Stage that can be used to reach high-energy orbits. Antares is topped by a 3.94-meter payload fairing.

Vehicle Description

Specifications

Type Antares
Height 40.5m
Diameter 3.90m
Launch Mass 282,000kg (110) - 296,000kg (130)
Mass to LEO 5,100kg (130)
Mass to GTO 1,800kg (132)
Mass to SSO 3,600kg (131)



First Stage

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Antares' first stage is designed by Yuzhnoye and built by Yuzhmash in the Ukraine. It is 27.6 meters long and 3.9 meters in diameter and contains Liquid Oxygen oxidizer and Rocket Propellant 1 fuel.

The Stage 1 assembly consists of the Stage 1 Core, the Main Engine System and the Flight Termination System. The Stage 1 Core is comprised of five bays and provides the structural support of the launch vehicle. It includes propellant tanks and associated plumbing, pressurization systems, instrumentation and avionics.

The Liquid Oxygen Tank Bay and Rocket Propellant 1 Bay consist of their corresponding propellant tanks which feature level sensors to measure propellant levels during fueling and in flight. This measurement system is used by the engine controllers to determine engine mixture ratio adjustments to minimize leftover residuals in the tanks. The RP-1 tank utilizes an aluminum waffle structure and it features a tunnel through its core to accommodate the LOX feedline. Routing the LOX feedline from the LOX tank above to the engines through the RP-1 tank increases packaging efficiency.

The LOX tank is manufactured from solid aluminum. It features Helium Pressurant Bottles which are submerged in the LOX tank for gas storage efficiency. Helium loading begins once the bottles are submerged in Liquid Oxygen in order to be chilled down to accommodate the Helium which is used to pressurize both, the RP-1 and LOX tanks. The pressurization system operates at a maximum pressure of 220 bar and supplies gas through a manifold of valves that are cycled open to regulate tank pressurization and propellant flow rates. Over pressurization is prevented by emergency relief valves.

The remaining three bays are the inter-tank bay, the inter-stage bay and the aft bay that includes the Main Engine System. The aft bay also contains the primary interface between the Antares launch vehicle and ground support equipment. Most of the mechanical, fluid and electrical interfaces of the launcher are located on its base.

The Main Engine System consists of two Aerojet AJ-26-62 engines. These are modified NK-33 engines that are being converted by Aerojet, importing the Russian engines and adding US electronics, making modifications to the fuel systems, modifying the engine for proper gimbaling control and removing engine harnesses. NK-33 engines were originally built for the massive Soviet N1 Moon Rocket in the 1960s and 70s and have since been in storage, not being used on any launcher.

The engine is a regeneratively cooled staged combustion engine with oxygen-rich preburners to drive the turbopumps. NK-33 provides a maximum sea level thrust of 1,630 Kilonewtons with length of 3.7 meters, an engine diameter of 2 meters and dry weight of 1,235kg. The engine can lift 137 times its weight and provides a vacuum impulse of 331s. Maximum vacuum thrust is 1,815kN. AJ-26 can be throttled from 56 to 108% of rated performance. The two AJ-26 engines of Antares are mounted on a thrust frame and each engine is equipped with independent Thrust Vector Control Systems for vehicle control during ascent. Engine gimbaling is controlled by a Moog hydraulic TVC system.

The two engines are fed by two separate LOX/RP-1 feed systems and have independent electrical hardware. The propellant inlets of the two engines are flexible to allow the engines to move relative to the core structure for Thrust Vector Control. The engine controllers are built by Orbital, but also incorporate engine sensors and propellant utilization systems provided by Yuzhnoye. Also part of the MES is the aft bay closure and the heat shield thermal blanket that protects aft bay components from the heat generated by the main engines.

At liftoff, the two engines generate a total thrust of 332,400 Kilograms. The first stage has a burn time of 235 seconds.

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First Stage

Type S1 Core Stage, Zenit Heritage
Bays 5
Dry Mass 18,700kg
Launch Mass 260,700kg
Diameter 3.90m
Length 27.6m
Fuel Rocket Propellant 1
Oxidizer Liquid Oxygen
Fuel Mass 64,740kg
Oxidizer Mass 177,260kg
Tank Pressurization Helium, up to 220bar
# Helium Bottles 8
Propulsion 2 x AJ-26-62 (Modified NK-33)
Throttling 56% - 108%
AJ-26 SL Thrust 1,510kN (100%) - 1,630kN (108%)
AJ-26 Vac Thrust 1,682kN (100%) - 1,815kN (108%)
Throttling 56% - 108% (Up to 135% achieved)
Impulse SL 297s
Impulse Vac 331s
Engine Length 3.71m
Engine Diameter 2.0m
Engine Dry Weight 1,235kg
Thrust to Weight 137
Chamber Pressure 145bar
Area Ratio 27
Ox. To Fuel Ratio 2.8
Burn Time 235sec
Attitude Control Hydraulic TVC for Yaw, Pitch & Roll
Stage Separation Hold Down Bolt Release

Second Stage

The second stage of the Antares launch vehicle is a solid rocket motor built by Alliant Techsystems, ATK. The first two mission of Antares, the Demo Flight and the Orb-D1 Demonstration Flight to ISS, will use a Castor 30A upper stage while the next two flights will feature the upgraded Castor 30B, flying on Orb-1 and Orb-2. Subsequent missions will use the stretched Castor XL to increase payload capability and allow an upgraded version of the Cygnus to carry more cargo to ISS.

Second Stage

Type Castor 30A
Length 3.51m
Diameter 2.34m
Dry Mass 1,220kg
Casing Mass 408kg
Ignition, Nozzle, TVC 340kg
Propellant Mass 12,815kg
Launch Mass 14,035kg
Propellant HTPB H8299
Avg. Thrust 259kN
Max Thrust 393kN
Chamber Pressure 53Bar
Specific Impulse 301s
Nozzle Diameter 1.21m
Expansion Ratio 65
Burn Time 136s
Vehicle Control Electromechanical TVC
Stage Separation Non-contaminating frangible ring
Attitude control Cold Gas ACS

Castor 30

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The Castor 30 solid rocket motor is based on ATK's Castor 120 which is a derivative of the first stage motor of the Peacekeeper missile that was in service from 1986 to 2005. Castor 120 was used on Lockheed Martin's Athena launch vehicles.

Castor 30A is 2.34 meters in diameter and 3.51 meters long. It has an empty weight of 1,220 Kilograms and can hold 12,815 Kilograms of propellant. The solid rocket motor uses a 408-Kilogram composite graphite/epoxy wound case.

Castor uses HTPB bound solid propellant. It is optimized for operation in vacuum conditions with an average thrust of 259 Kilonewtons, peaking up to 393 Kilonewtons. The second stage provides a specific impulse of 301 seconds. Castor 30A burns for 136 seconds. Ignition is accomplished with an IVB Ignitor. A flex seal design at the throat of the SRM allows nozzle motion during flight for Thrust Vector Control. The nozzle can be gimbaled as part of a Electromechanical Thrust Vector Control System. A 65:1 throat-to-exit-plane-ratio model provides the second stage performance for the initial Antares missions ahead of second stage upgrades.

The Castor 30B rocket motor is a higher-performance version of the Castor 30A with an increase in Isp of under 5 seconds. Its overall length is 4.17 meters and it has a diameter of 2.34 meters. The motor features and extended nozzle with a 76:1 expansion ratio and a diameter of 1.63 meters. The 30B version provides an average thrust of 293.4 Kilonewtons and a maximum thrust of 396.3 Kilonewtons. Castor 30B has a liftoff weight of 13,970 Kilograms and burns for 127 seconds.

Castor 30

Type Castor 30B
Length 4.17m
Diameter 2.34m
Propellant Mass 12,887kg
Launch Mass 13,970kg
Propellant HTPB H8299
Avg. Thrust 293.4kN
Max Thrust 396.3kN
Chamber Pressure 53Bar
Specific Impulse ~304s
Nozzle Diameter 1.63m
Expansion Ratio 76
Burn Time 127s
Vehicle Control Electromechanical TVC
Stage Separation Non-contaminating frangible ring
Attitude control Cold Gas ACS

Castor 30 XL

The Castor 30 XL is a stretched version of the Castor 30A solid rocket motor. It also uses a composite graphite/epoxy wound case and HTPB bound propellant. Castor 30 XL is 2.34 meters in diameter and 5.99 meters long. It has a liftoff mass of about 26,300 Kilograms. The stage delivers an average thrust of over 300kN peaking at 395kN.

Castor 30 XL uses a 2.4-meter long nozzle with a high expansion ratio of 56:1 and submerged design. A dual density exit cone improves performance for operation at high altitudes. The nozzle design features a number of changes to the Castor 30A nozzle such as a modified Propulsion Application Program flex bearing with a 3.5-degree maximum design vector angle. The Castor 30 XL also uses an Electromechanical Thrust Vector Control System that is identical to that of Castor 30A.

Castor 30 XL

Type Castor 30 XL
Length 5.99m
Diameter 2.36m
Launch Mass ~26,300kg
Propellant HTPB H8299
Nozzle Length 2.4m
Expansion Ratio 56
Burn Time 156s
Vehicle Control Electromechanical TVC
Stage Separation Non-contaminating frangible ring
Attitude Control Cold Gas ACS

Stage 2 Avionics & Guidance System

The Antares avionics module is based on Orbital's Modular Avionics Control Hardware (MACH) that will provide power transfer, data acquisition, booster interfaces, and ordnance initiation. Most avionics are located in the avionics ring mounted on the second stage. Antares uses a three-axis autopilot that utilizes Proportional-Integral-Derivative control. The first stage flies a pre-programmed attitude profile based on trajectory design and optimization while the second stage adjusts its flight profile dynamically to achieve a pre-programmed set of orbital parameters. It uses energy management to place the vehicle into its target orbit.

Second Stage Attitude Control System

During the second stage burn, a combination of TVC and ACS is used. Antares features a three-axis cold gas attitude control system on its second stage to provide orientation capability during coast phases, for spacecraft separation and subsequent collision avoidance maneuvers.

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Optional Third Stage

Antares can fly with a third stage to increase payload capability and provide accurate insertion capabilities for high-energy insertions.

Optional Third Stage: Star-48BV

Type Star 48BV
Launch Mass 2,164.5kg
Diameter 1.24m
Length 2.08m
Propellant TP-H-3340
Propellant Mass 2,010.0kg
Casing Mass 58.3kg
Case Material Titanium
Nozzle Mass 52.6kg
Avg Thrust 68.6kN
Max Thrust 77.8kN
Isp 288s
Throat Diameter 0.1011m
Nozzle Diameter 0.7475m
Chamber Pressure 39.9bar (Avg) - 42.6bar (Max)
Expansion Ratio 54.8
Burn Time 84.1s
Ignition Delay 0.100s
Attitude Control TVC +/-4°

Optional Third Stage: BTS

Type BTS
Fuel Monomethylhydrazine
Oxidizer Nitrogen Tretroxide
Fuel Mass 358kg
Oxidizer Mass 322kg
Propulsion 3 x BT-4
BT-4 Thrust 450N
Total Thrust 1,350N
Engine Mass 4kg
Engine Length 0.65m

ATK Star 48BV

The ATK Star 48BV is a solid-propellant upper stage that uses the flight proven Star 48B and adds Thrust Vector Capability (V). Star 48 was introduced in 1982 and has been used on a variety of spacecraft. Star 48B was spin stabilized and had smaller a performance than the TVC capable version.

The upper stage features a 1.24-meter diameter titanium casing holding a total of 2,010 Kilograms of solid propellant. It is 2.08 meters in length and has a launch mass of 2,165 Kilograms. It operates at an average thrust of 68.6 Kilonewtons with peak thrust reaching 77.8kN. Star-48BV features the longer of two available nozzles for the conventional Star 48. The upper stage features an electromechanically actuated flexseal nozzle Thrust Vector Control System with a maximum nozzle gimbal of four degrees. Star-48 burns for 84 seconds and is suitable for payloads that are inserted into high-energy trajectories.

Bi-Propellant Third Stage

The Bi-Propellant Third Stage, BTS, is provided by Orbital Sciences and is based on Orbital's GEOStar satellite bus that is used for Geostationary Satellites. The Upper Stage features a Helium-regulated bi-propellant system with Monomethylhydrazine fuel and Nitrogen Tetroxide oxidizer.

The propellants are stored in spherical tanks. A total of 358 Kilograms of MMH and 322 Kilograms of NTO can be carried by the vehicle. The propulsion system consists of three IHI BT-4 engines. BT-4 was developed by IHI Aerospace, Japan, and has a dry mass of 4 kilograms and a length of 0.65 meters. The engine provides 450 Newtons of Thrust.

The BTS can perform precise insertions and multiple engine burns for orbit circularization. Sun Synchronous Missions of Antares would typically use this upper stage.

Payload Fairing

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The Payload Fairing is positioned on top of the launch vehicle and its integrated Payload. It protects the spacecraft against aerodynamic, thermal and acoustic environments that the vehicle experiences during atmospheric flight. When the launcher has left the atmosphere, the fairing is jettisoned by pyrotechnical initiated systems. Separating the fairing as early as possible increases launcher performance.

Antares jettisons its fairing during the Coast Phase between the first and second stage burn during a typical LEO Flight.

The fairing is 3.94 meters in diameter and 9.87 meters long. It consists of two composite shell halves, a low-shock frangible rail and ring separation system, and an actuator/hinge fairing jettison system. The fairing structure is a aluminum honeycomb core covered by layers of graphic epoxy composite. The two fairing halves are joined by a frangible rail joint and the PLF is connected to the second stage using a ring-shaped frangible joint.

Ordnances and a clean-separation frangible joint with a confined sealed stainless steel tube is used to fracture aluminum extrusions on the ring and rail. Disconnecting the ring and rail allows each half of the fairing to rotate on hinges installed on the second stage. A cold gas generation system is utilized to drive pistons that force the fairing halves to open.

Payload Adapters

Payload Adapters interface with the launch vehicle and the payload and are the only attachment point of the payload on the Launcher. They provide equipment needed for spacecraft separation and connections for communications between the Upper Stage and the Payload.

Orbital Sciences offers a number standard payload adapters to install spacecraft on the Antares launch vehicle. Payload Adapters are provided by RUAG Space and include the 1194VS, 1666 and 937 payload systems that provide accommodations for a number of spacecraft and feature low-shock separation techniques.

Antares LEO Flight Profile

Antares lifts off from its launch pad two seconds after the AJ-26 engines of the first stage are ignited to allow some time for them to achieve full thrust and monitor their ignition performance. After a short vertical ascent, Antares performs a roll & pitch maneuver to align itself with its pre-planned ascent trajectory.

The first stage burns for 235 seconds and separates after a brief, 5-second post-burn coast. Stage 1 separation occurs at an altitude of 109 Kilometers and a velocity of 4,547m/s. At that point, the stack enters a 100-second coast period to get close to apogee for the second stage burn. After 100 seconds of coasting, the Payload Fairing is jettisoned at an altitude of 184 Kilometers. Ten seconds later, the second stage begins its engine burn for orbital insertion and circularization. Stage 2 shutdown occurs about 471 seconds into the flight at an altitude of 205 Kilometers and a velocity of 7,521m/s. Payload separation occurs after 120 seconds of maneuvering by the second stage attitude control system. The typical Cygnus insertion orbit is 275 by 250 Kilometers at an inclination of 51.66 degrees.

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’We've Got a Satellite!’

’We’ve Got a Satellite!’

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With the possible exception of Columbia and the very first space shuttle mission, few orbiters had as dramatic and exciting a maiden voyage as Endeavour. On STS-49, she provided a reliable stage for the longest EVA in history and the first three-man EVA in history. Photo Credit: NASA

“Ready. Ready. Grab!

The words of Rick Hieb echoed through the silent Mission Control Center (MCC) at the Johnson Space Center (JSC) in Houston, Texas.

The view through Space Shuttle Endeavour’s aft flight deck windows on the evening of 13 May 1992 was quite different from anything ever seen before. Not only was this the maiden voyage of NASA’s newest orbiter—a vehicle which, but for the loss of Challenger, might have remained a set of structural spares—but it also involved the first (and only) EVA with as many as three people. This mission, STS-49, commanded by Chief Astronaut Dan Brandenstein, had long been anticipated to be the most visible shuttle flight of 1992, but it demonstrated that human space flight retains the ability to deliver unexpected surprise. When the crew was announced, their mandate was to retrieve the Intelsat 603 telecommunications satellite, delivered into an improper orbit by a Commercial Titan III booster in March 1990. Spacewalkers Hieb and Pierre Thuot would venture into Endeavour’s payload bay to attach a new rocket motor, after which Intelsat 603 would be boosted into its 22,300-mile (35,900-km) geosynchronous orbit, ahead of its pivotal role in covering the 1992 Summer Olympics in Barcelona.

After the Intelsat activities, a further two spacewalks—the first with Kathy Thornton and Tom Akers, the second with Thuot and Hieb—would rehearse Space Station Freedom construction techniques. Thornton’s inclusion made her only the third woman, after Russia’s Svetlana Savitskaya and NASA’s Kathy Sullivan, to perform an EVA. It was a role for which she had previously trained in preparation for her first shuttle mission, STS-33 in November 1989. “I absolutely insisted that she be the EVA person,” STS-33 Commander Fred Gregory recalled in his NASA Oral History, “over great protest. If we had not insisted, probably a person of her size would never have done something like this. Kathy [Sullivan] was a larger woman who could fit into the suits, but Kathy Thornton was not, so we really had to force the issue.” Doubtless, Dan Brandenstein was in full agreement that Thornton, nicknamed “K.T.”, was the most appropriate choice. She would go on to fly as part of the EVA team which first serviced the Hubble Space Telescope (HST) in December 1993.

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The other three astronauts involved in the STS-49 EVAs were male. Rick Hieb was already in training to fly STS-39 at the time the Intelsat 603 crew was assembled in December 1990, and Tom Akers had returned only weeks earlier from the Ulysses deployment mission, STS-41. The man in charge of the team—designated “EV1” and wearing red stripes around the legs of his pure-white space suit for identification—was Pierre Thuot. When he flew STS-36 in the spring of 1990, Thuot became the first of his class to be assigned a mission and the first to actually fly.

If everything ran as timelined, STS-49 would thus be the first shuttle flight to feature as many as three spacewalks and the first to include two teams of spacewalkers; both of which were critical prerequisites if NASA was to execute as many as five EVAs per mission to service the Hubble Space Telescope (HST) and build Space Station Freedom. On the face of it, retrieving and repairing Intelsat 603, for all its drama, offered something of a backward glance to the shuttle’s pre-Challenger heyday and was unusual, for in the wake of the disaster it had been mandated that the reusable orbiters would henceforth not be used for commercial missions. STS-49 was thus the last of its kind. At the same time, as Space Shuttle Program Director Bob Crippen explained in June 1990, it offered “an opportunity for expanding our experience base in the planning, training and performance of EVA” by “helping preparations for Freedom.”

Others agreed that such a mission was useful for other purposes. It was “a throwback to the good old days,” said Endeavour’s first processing manager, John Talone, “when we used to go out and do these kinds of things.” Added NASA Associate Administrator for Space Flight, former astronaut Bill Lenoir: “It’s a mission we wanted to do. It gave me the opportunity to have real work that really mattered; that was going to get measured, where we either succeeded or failed.”

In the weeks and months following the botched delivery of Intelsat 603 to orbit in March 1990, prime contractor Hughes entered into a contract with NASA, worth in excess of $90 million, for the shuttle to reboost the satellite. Two possible scenarios quickly gained prominence: either to attach a new perigee kick motor or retrieve Intelsat and bring it back to Earth for refurbishment. Concerns about the extent to which the satellite’s surfaces might degrade over two years were allayed by the test flight of several solar array sample “coupons,” attached to Discovery’s Remote Manipulator System (RMS) mechanical arm during the STS-41 mission in October 1990. These were exposed to the harsh atomic oxygen environment for a minimum of 23 hours, with few ill-effects. Two months later, in December, the STS-49 crew was named to conduct the audacious salvage.

Dan Brandenstein found himself in command of the first flight of a new shuttle and a rendezvous and retrieval mission with EVAs which promised to be filled with drama. “One of my first concerns when we first got assigned and started working with Hughes on the mission,” he told the NASA oral historian, “was if we try and grab it, if we bump it, is it going to go out of whack and float away? Part of the requirements from the customer were that we didn’t touch any sensitive area, which left you a very small ring that … had a limited accessibility and that was supposed to the way we grabbed it.”

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Armed with the capture bar mechanism, Pierre Thuot provides a measure of scale of the enormous size of Intelsat 603. Photo Credit: NASA

The mechanism by which Thuot and Hieb would grab Intelsat was a so-called “capture bar,” designed and built by engineers in the Crew and Thermal Systems Division at NASA’s Johnson Space Center (JSC) in Houston, Texas. Weighing 160 pounds (73 kg), it measured 15 feet (4.6 meters) long by about 3.3 feet (1 meter) wide and included detachable beam extensions and a steering wheel. As Thuot rode on the end of Endeavour’s Remote Manipulator System (RMS) mechanical arm, he would be positioned close to the base of Intelsat 603 and after grappling it would lower it delicately into a Hughes-built cradle assembly. “There was a lot of analysis done,” continued Brandenstein, “and we were assured that because it was spinning slightly and it had a lot of mass, we could bump it and it would stay pretty much in place and wasn’t going to be a problem.” Throughout 1991 Thuot and Hieb trained underwater and on the air-bearing table, to such an extent that they could follow the procedure with their eyes closed.

More than two decades ago, on 7 May 1992, the last virgin space shuttle speared for the heavens. During the next couple of days Intelsat controllers maneuvered their satellite into a “control box,” some six degrees of arc of the shuttle’s orbit. These maneuvers also served to reduce Intelsat’s rotation from 10.5 to around 0.65 rpm. By the 10th, as they approached to within 8 miles (13 km) of the satellite, Thuot and Hieb completed their procedures of suiting-up and were assisted into the airlock by crewmate Akers. Shortly thereafter, at 4:25 p.m. EDT, they opened Endeavour’s outer hatch into the payload bay—then in the pitch black of orbital darkness—and Thuot fastened himself into a foot restraint on the end of the RMS, deftly operated by veteran astronaut Bruce Melnick. Drawing closer toward the satellite, Thuot extended the capture bar into position, but the latches failed to latch.

He tried again, without success.

A third attempt was similarly fruitless.

From his station on Endeavour’s aft flight deck, Brandenstein noticed that Intelsat 603 was beginning to oscillate and drift somewhat, “so I got in my chase-it mode, because I had to keep him aligned.” When Thuot’s third attempt failed, Brandenstein had used a “tremendous” amount of propellant and instinctively knew that the chances of success were slim at best. The RMS exacerbated the difficulty, because its joints were being driven into positions which they could not support. “We decided, though consultations with the ground, to get out of there and try another day,” Brandenstein recollected. “That was a pretty low point, because when we left, it had a pretty good rate. We thought we’d lost this $150 million satellite … and Pierre was particularly depressed because, obviously, he thought it was his fault.”

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In the first, and so far only, three-person EVA, astronauts Rick Hieb, Tom Akers, and Pierre Thuot manhandle Intelsat 603 into Endeavour’s payload bay for the attachment of a new rocket motor. Photo Credit: NASA

Thuot and Hieb returned inside Endeavour after three hours and 43 minutes, and later that evening Hughes engineers confirmed that they had managed to stabilize Intelsat. Next day, at 4:30 p.m. EDT on 11 May, the spacewalkers were back outside for a second attempt. “Instead of doing it at night, we were going to wait and do it in daylight,” Brandenstein said. “We decided we weren’t going to even make an attempt until everything was just perfect. Pierre went in and the rotation slowed down.” From Hieb’s position, it looked as if Thuot had completed the capture, but, alas, the satellite again began to oscillate. The astronaut’s alignment was unquestionably correct, but the capture bar refused to seat itself properly and Intelsat wobbled. A few weeks after the mission, Thuot explained to this author that the satellite “was much more dynamic than our training had led us to believe.”

As the disappointed spacewalkers returned inside the cabin for the second time—this time after 5.5 hours—they at least knew that the Hughes engineers could regain control of Intelsat 603 for another attempt. However, although propellant reserves allowed for it, three separate rendezvous on a single shuttle mission had never been attempted, and Brandenstein recommended a day off to plan for the third attempt. In an interview for the Smithsonian, Rick Hieb remembered that the evening of the 11th was a sombre time. At one point, Kevin Chilton, the STS-49 pilot and the only “rookie” member of the crew, joined Hieb on the flight deck, and the pair entered an impromptu brainstorming session. It was a session that would mark a significant turnaround in the fortunes of a mission which seemed snake-bitten.

As Hieb and Chilton talked, other members of the crew floated upstairs to join them. The main concern was where to manually grab Intelsat. The top of the satellite, where the delicate antennas were located, was not ideal, and it was Bruce Melnick who suggested an EVA with not two spacewalkers, but three. No excursion in history had ever involved more than two members, partly due to safety concerns and partly because of the sheer practicality of getting three people into the tiny airlock. On the other hand, Endeavour carried four suits for Thuot, Hieb, Thornton, and Akers, so in theory it was a possibility.

“When Bruce said that,” recalled Hieb, “a big mental switch flipped over, at least for me. In my mind, having a third set of hands out there meant that we would be successful, although we weren’t yet sure how.”

Mission Control knew that the astronauts were still awake, because Endeavour’s monitors had not been turned off. At length, the crew turned them off and continued talking in the dark, but eventually called the ground with Melnick’s idea. Years later, Brandenstein remembered that it was Chilton who sketched out the practicalities of the three-person EVA scenario and held it in front of the television camera to allow mission controllers to see it. “The big choke point,” Brandenstein said, “was can you put three people in the airlock to get them outside?” In the Houston water tank, fellow astronauts Story Musgrave, Jim Voss, and Michael “Rich” Clifford donned suits and demonstrated the techniques and geometries involved in setting themselves up to accomplish the feat.

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Proudly demonstrative of her nautical and exploratory heritage through Captain James Cook’s vessel, and bearing the colors of the elementary and secondary schools which named her, Endeavour’s first patch is bordered by the names of her first seven astronauts: Commander Dan Brandenstein, Pilot Kevin Chilton and Mission Specialists Rick Hieb, Bruce Melnick, Pierre Thuot, Kathy “K.T.” Thornton, and Tom Akers. Image Credit: NASA

Their consensus: It was doable. On the evening of 12 May, Capcom Charles “Sam” Gemar radioed Mission Control’s approval to the crew.

Late on 13 May, the third attempt got underway. Truss members belonging to the Assembly of Station by EVA Methods (ASEM)—a Space Station Freedom demonstration payload, to be used during EVA tests later in the mission—were removed and arranged into a triangular structure for Thuot, Hieb, and Akers to anchor their feet. Brandenstein positioned the orbiter directly beneath Intelsat 603, and controllers verified that its surface temperatures would not exceed the 160 degrees Celsius (320 degrees Fahrenheit) touch limit of the astronauts’ gloves. With Hieb close to the starboard payload bay wall, Akers in the center, attached to an ASEM strut, and Thuot on the end of the RMS on the port side, the astronauts could do little but watch as Endeavour drew closer. They studied its slow rotation for about 15 minutes, until, on Hieb’s call, they moved in for the capture.

All at once, Thuot spotted a slight wobble. He called the attempt off.

Shortly thereafter, they tried again. This time, at last, the three men grabbed the satellite and held it firmly. The time was 7:55 p.m. EDT. “I actually thought the other two guys had stopped it from rotating,” Thuot said later, “so little force had I applied. Very gently, the thing came to a stop.” From the flight deck, Dan Brandenstein asked them if they had a good grasp. On Thuot’s response in the affirmative, the commander was able to advise ground controllers, with more than a hint of relief: “Houston, I think we’ve got a satellite!”

With Intelsat snared, the astronauts removed the steering wheel and installed an extension to the capture bar, which Melnick grappled using the RMS. The satellite was then positioned above its 23,000-pound (10,430-kg) Orbus-21 solid-fueled perigee kick motor, which sat vertically in its cradle. After closing four docking clamps to secure the pair, and attaching two electrical umbilicals between Intelsat and the motor itself, the spacewalkers set a pair of deployment timers and retreated to Endeavour’s airlock. Meanwhile, Kathy Thornton prepared to activate the springs to deploy the payload. At first, it did not move. “They had made a change in the wiring of the deploy system,” recalled Brandenstein, “and the change never made it through the process [and] never got into the checklist. Fortunately, somebody in Mission Control apparently knew about it. They just quick called up a different switch sequence and she did that sequence and it went.” Deployment occurred at 12:53 a.m. EDT on 14 May, and the satellite vacated the payload bay. Less than an hour later, the three spacewalkers repressurized the airlock and returned inside the cabin.

Speaking a decade or more after the flight, Dan Brandenstein regarded those few days of STS-49 as “one of those missions from hell,” and for newly-appointed NASA Administrator Dan Goldin it was truly “a baptism of fire.” Nevertheless, at 1:25 p.m. EDT on the following day, 15 May, Intelsat 603’s new motor ignited perfectly, and it was on-station in geosynchronous orbit by the 21st. As well as becoming the first shuttle crew to accomplish as many as three EVAs in a single mission—a record which they would break with a fourth excursion—the triumphant three-man spacewalk established itself as the longest in history. Their eight hours and 29 minutes outside would remain unbroken until March 2001. By now, the difficulties had prompted the Mission Management Team (MMT) to extend STS-49 by 48 hours from its planned seven-day duration. On 14 May, a record-breaking fourth EVA got underway when Akers and Kathy Thornton ventured outside for the ASEM station tests.

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Kathy “K.T.” Thornton and Tom Akers participate in the mission’s record-breaking fourth EVA to perform Space Station evaluations. Photo Credit: NASA

Originally scheduled to involve two EVAs—one by Thornton and Akers and the second by Thuot and Hieb—the Intelsat 603 retrieval forced the cancelation of one spacewalk.

Activities included the construction of a pyramid-shaped truss, the unberthing of a Mission-Peculiar Equipment Support Structure (MPESS)—maneuverd by the RMS—and efforts to evaluate the ability of spacewalkers to work at positions “above” and “forward” of the payload bay, including “over the nose” of the shuttle. The MPESS contained two node boxes for the pyramid, a releasable grapple fixture and interface plate, and a truss leg and strut dispenser. Five crew rescue techniques were to be trialed, including a lasso-like “astro-rope,” a seven-section telescoping pole, and a hand-held propulsive device. The latter, according to NASA’s STS-49 press kit, was “a redesigned hand-held maneuvering unit from the Skylab program,” in which pressurized nitrogen jets were employed as thrusters.

During their seven hours and 43 minutes in the payload bay, Thornton and Akers completed the construction and disassembly of the ASEM attachment fixture, tested the propulsive device, affixed six of eight legs onto the MPESS and, unexpectedly, were called upon to manually stow Endeavour’s Ku-band antenna, which had experienced a positioning motor failure. According to NASA’s post-mission report, this EVA was planned to be RMS-intensive, although the mechanical arm was used to accomplish only a single ASEM task and the spacewalkers’ timeline was further impacted by the Ku-band activity.

Returning inside Endeavour’s airlock after the excursion, the astronauts of STS-49 could now boast four EVAs—lasting a cumulative of 25 hours and 27 minutes—which had snatched success from the fangs of defeat. The physical appearance of the four spacewalkers in their snow-white suits was also quite distinct from previous missions, all of which had featured no more than two members. In order to distinguish them, Thuot (designated “EV1”) wore red stripes around his suit legs, whilst Hieb (EV2) wore a pure-white suit, and, for the first time, Thornton (EV3) wore dashed stripes around her suit legs and Akers (EV4) wore red diagonal hatches around his suit legs. In spite of the remarkable achievement of performing four back-to-back EVAs on a single mission, only relatively minor glitches plagued the spacewalkers—a failed joint on one of the portable foot restraints, a loud noise over the headsets when power tools were being used, and a battery problem, amongst others—and their suits held up exceptionally well.

In the aftermath of STS-49, the crew themselves would highlight the fact that their mission raised awareness of the need for more EVA experience in the years before the start of construction of Space Station Freedom. At one stage, in the late 1980s, as many as four EVAs per week were envisaged—an astonishing estimate which NASA Administrator Dick Truly deemed totally unacceptable. Yet as the plans for the station matured, it was obvious that the construction process would be EVA-intensive and that required different ways of working and training. “We have to take a good look at the time it takes to do a job,” Brandenstein said. “We need better ways to train so that the learning curve isn’t quite so steep.” Pierre Thuot added that the Intelsat 603 retrieval task was something that the crew “couldn’t train for fully.”

After their return from STS-49, Kathy Thornton and Tom Akers—who would go on to service the Hubble Space Telescope (HST) together in December 1993—took an active role in developing new EVA methods in the Weightless Environment Training Facility (WET-F) in Houston. “Even in the tank, you still have the resistance of water,” Dan Brandenstein recalled, “so you can kick your feet and swim. In zero gravity, we’ve got movies of Tom going to that instinct. You can see him kicking his legs and nothing’s happening. Also, if you move something in the water, as soon as you stop moving it, the reason is the water stops it. But in zero gravity, you start moving something and it just keeps moving until you come back on it. They made some significant chances in the tank training procedures.” The first flight of Endeavour’s career had gone spectacularly well and had played a significant role in shaping the missions—and the assembly of the space station—which would follow.
 

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