J-2 Facts
Length 11ft 1in. Width 6ft 8 1/2in. Thrust (altitude) 200,000 lb Specific Impulse (nominal) 426 sec Rated Run duration 500 sec Flowrate (oxidizer) 392 lb/sec (2490 gpm) Flowrate (fuel) 78 lb/sec (7960 gpm) Mixture ratio 5:1 oxidizer to fuel Nominal Chamber pressure 632 psi Weight, dry flight config. 3480 lb Expansion Area ratio 27.5:1 J-2 ENGINE DESCRIPTION
The J-2 engine is a 200,000 pound thrust, high performance, upper stage, propulsion system, utilizing liquid hydrogen and liquid oxygen propellants, and incorporating a built-in capability for restart in flight.
Major systems of the J-2 engine include a thrust chamber and gimbal system, propellant feed system, gas generator and exhaust system, electrical and pneumatic control system, start tank assembly system, and flight instrumentation system.
Thrust Chamber and Gimbal System
The J-2 engine thrust chamber serves as a mount for all engine components. It is composed of the following subassemblies: thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark igniter. Thrust is transmitted through the gimbal mounted on the thrust chamber assembly dome to the vehicle thrust frame structure. The thrust chamber and injector receives the propellants from the turbopumps under pressure, mixes the propellants, and burns them to impart a high velocity to the expelled combustion gases to produce thrust for vehicle propulsion.
THRUST CHAMBER
The thrust chamber is constructed of stainless steel tubes of 0.012-inch wall thickness. Tubes with thin walls are required for heat transfer purposes. The thrust chamber tubes are stacked longitudinally and furnace brazed to form a single unit. The chamber is bell shaped with 27.5 to 1 expansion area ratio for efficient operation at altitude, and is regeneratively cooled by the fuel.
Fuel enters from a manifold to which it was delivered at a pressure of more than 1000 psi. It makes a one-half pass downward through 180 tubes and a full pass up through 360 tubes to cool the chamber.
DOME
The injector and dome assembly is located at the top of the thrust chamber. The dome manifolds the liquid oxygen and serves as a mount for the gimbal bearing and the augmented spark igniter.
INJECTOR
The thrust chamber injector is concentric orificed and porous faced. The purpose of the thrust chamber injector is to atomize and mix propellants in a manner to produce the most efficient combustion. Hollow oxidizer posts are machined to form an integral part of the injector. Fuel nozzles are threaded and installed over the oxidizer posts.
The injector face is formed from porous stainless steel and is welded at its periphery to the injector body. Each fuel nozzle is swaged to the face of the injector.
The injector receives liquid oxygen through the dome manifold and injects it through the oxidizer posts into the combustion area of the thrust chamber.
The fuel is received from the upper fuel manifold in the thrust chamber and injected through the fuel orifices which are concentric with the oxidizer orifices. The propellants are injected uniformly to ensure satisfactory combustion.
GIMBAL
The gimbal is a compact, highly-loaded (20,000 psi), universal joint consisting of a spherical, socket-type bearing with a Teflon-fiberglass composition coating that provides a dry, low-friction bearing surface. It also includes a lateral adjustment device for aligning the chamber with the vehicle.
The gimbal transmits the thrust from the injector assembly to the vehicle thrust structure and provides a pivot bearing for deflection of the thrust vector, thus providing flight attitude control of the vehicle. The gimbal is mounted on the top of the injector assembly.
AUGMENTED SPARK IGNITER
The augmented spark igniter (ASI) is mounted in the injector. It receives the initial flow of oxidizer and fuel, which is ignited by two spark plugs mounted in the side of the igniter chamber. When engine start is initiated, the spark exciters energize the spark plugs. Simultaneously, the control system starts the flow of oxidizer and fuel to the spark igniter. As the oxidizer and fuel enter the main combustion chamber, they are mixed and ignite.
Mounted in the ASI is an ignition monitor which indicates that proper ignition has taken place. The ASI operates continuously during entire engine firing, is uncooled, and is capable of multiple reignitions under all environmental conditions.
Propellant Feed System
The propellant feed system consists of separate oxidizer and fuel turbopumps, main fuel valve, main oxidizer valve, propellant utilization valve, oxidizer and fuel flowmeters, fuel and oxidizer bleed valves, and interconnecting lines.
OXIDIZER TURBOPUMP
The oxidizer turbopump is mounted on the thrust chamber diametrically opposite the fuel turbopump. It is a single- stage centrifugal pump with direct turbine drive. The oxidizer turbopump increases the pressure of the liquid oxygen and pumps it through high-pressure ducts to the thrust chamber at a rate of 2,475 gallons per minute. The pump operates at 8,000 rpm at a discharge pressure of 930 psia and develops about 1,700 brake horsepower. The pump and its two turbine wheels are mounted on a common shaft.
Power for operating the oxidizer turbopump is provided by a high-speed, two-stage turbine which is driven by the exhaust gases from a bi-liquid gas generator. The turbines of the oxidizer and fuel turbopumps are connected in series by exhaust ducting that directs the discharged exhaust gas from the fuel turbopump turbine to the inlet of the oxidizer turbopump turbine manifold. One static and two dynamic seals in series prevent the turbopump oxidizer fluid and turbine gas from mixing.
The turbopump operates in this manner: hot gas enters the nozzle and, in turn, the first-stage turbine wheel. After passing through the first-stage turbine wheel, the gas is redirected by the stator blades and enters the second-stage turbine wheel. The gas then leaves the turbine through exhaust ducting, passes through the heat exchanger, and exits through the thrust chamber. Power from the turbine is transmitted to the inducer and impeller by the pump shaft. The velocity of the liquid oxygen is increased through the inducer and impeller. As the liquid oxygen enters the outlet volute, velocity is converted to pressure and the liquid oxygen is discharged into the outlet duct at high pressure.
FUEL TURBOPUMP
The fuel turbopump, also mounted on the thrust chamber, is a turbine-driven, axial flow pumping unit consisting of an inducer, a seven-stage rotor, and a stator assembly. It is a high-speed pump operating at 26,000 rpm, and is designed to increase hydrogen pressure to 1,140 psia through high-pressure ducting to the thrust chamber. It develops 6,750 brake horsepower.
Power for operating the turbopump is provided by the high-speed, two-stage turbine. Hot gas from the gas generator is routed to the turbine inlet manifold which distributes the gas to the nozzle where it is expanded and directed at a high velocity into the first-stage turbine wheel.
After passing through the first-stage turbine wheel, the gas is redirected through the stator blades and enters the second-stage turbine wheel. The gas leaves the turbine through the exhaust ducting. Three dynamic seals in series prevent the pump fluid and turbine gas from mixing. Power from the turbine is transmitted to the pump by means of a one-piece shaft.
Bearings in the liquid hydrogen and liquid oxygen turbopumps are lubricated by the fluid being pumped, as the extremely low operating temperature of the engine precludes use of lubricants or other fluids.
MAIN OXIDIZER VALVE
The main oxidizer valve is a butterfly-type valve, spring-loaded to the closed position, pneumatically operated to the open position, and pneumatically assisted to the closed position. It is mounted between the oxidizer high-pressure duct from the oxidizer turbopump and the oxidizer inlet on the thrust chamber assembly.
The main oxidizer valve controls the flow of oxidizer to the thrust chamber. As the oxidizer pressure in the high-pressure duct upstream of the main oxidizer valve builds up, pressure is sensed at the main oxidizer pressure-actuated sequence control valve. At a set pressure the control valve opens, allowing flow of pneumatic pressure to the opening control port of the main oxidizer valve second-stage actuator to fully open the valve and allow flow of oxidizer to the thrust chamber.
Pneumatic pressure from the normally closed port of the mainstage control solenoid valve is routed to both the first and second stage opening control ports of the main oxidizer valve. Controlled opening of the valve is accomplished by closing control pressure being bled from the valve through a restrictor check valve.
MAIN FUEL VALVE
The main fuel valve is a butterfly-type valve, spring-loaded to the closed position, pneumatically operated to the open position, and pneumatically assisted to the closed position. It is mounted between the fuel high-pressure duct from the fuel turbopump and the fuel inlet manifold of the thrust chamber assembly. The main fuel valve controls the flow of fuel to the thrust chamber. Pressure from the ignition stage control valve on the pneumatic control package opens the valve during engine start. As the gate starts to open, it allows fuel to flow to the fuel inlet manifold.
PROPELLANT UTILIZATION VALVE
The propellant utilization valve is an electrically operated, two-phase, motor-driven, oxidizer transfer valve and is located at the oxidizer turbopump outlet volute. The propellant utilization valve ensures the simultaneous exhaustion of the contents of the propellant tanks. During engine operation, propellant level sensing devices in the vehicle propellant tanks control the valve gate position for adjusting the oxidizer flow to ensure simultaneous exhaustion of fuel and oxidizer.
The propellant utilization (PU) valve and its servomotor are supplied with the engine. A position feedback potentiometer is also supplied as a part of the PU valve assembly. The PU valve assembly and a S-IVB stage or a facility-mounted control system make up the propellant utilization system.
OXIDIZER AND FUEL FLOWMETERS
The oxidizer and fuel flowmeters are identical helical-vaned, rotor-type flowmeters. They are located in the oxidizer and fuel high-pressure ducts. The flowmeters measure propellant flow rates in the high-pressure propellant ducts. The four-vane rotor in the hydrogen system produces four electrical impulses per revolution and turns approximately 3600 revolutions per minute at nominal flow. The six-vane rotor in the liquid oxygen system produces six electrical impulses per revolution and turns at approximately 2400 revolutions per minute at nominal low.
PROPELLANT BLEED VALVE
The propellant bleed valves used in both the oxidizer and fuel systems are poppet-type which are spring loaded to the open position and pressure actuated to the closed position. Both propellant bleed valves are mounted to the bootstrap lines adjacent to their respective turbopump discharge flanges.
The valves allow propellant to circulate in the lines to achieve proper operating temperature prior to engine start. The bleed valves are engine controlled. At engine start, a helium control solenoid valve in the pneumatic control package is energized allowing pneumatic pressure to close the bleed valves, which remain closed during engine operation.
Gas Generator Exhaust System
This system consists of the gas generator, gas generator control valve, turbine exhaust system and exhaust manifold, heat exchanger, and oxidizer turbine by pass valve.
GAS GENERATOR
The gas generator produces hot gases to drive the oxidizer and fuel turbines and consists of a combustor containing two spark plugs, a control valve containing oxidizer and fuel ports, and an injector assembly.
When engine start is initiated, the spark exciters in the electrical control package are energized providing energy to the spark plugs in the gas generator combustor. Propellants flow through the control valve to the injector assembly and into the combustor where they are mixed and burned. The resulting hot gases pass through the combustor outlet and are directed to the fuel turbine and then to the oxidizer turbine.
GAS GENERATOR CONTROL VALVE
The gas generator control valve is a pneumatically operated poppet type that is spring-loaded to the closed position. The oxidizer and fuel poppets are mechanically linked by an actuator. The gas generator control valve controls the flow of propellants through the gas generator injector.
When the mainstage signal is received, pneumatic pressure is applied against the gas generator control valve actuator assembly which moves the piston and opens the fuel poppet. During the fuel poppet opening, an actuator contacts the piston that opens the oxidizer poppet. As the opening pneumatic pressure decays, spring loads close the poppets.
TURBINE EXHAUST SYSTEM
The turbine exhaust ducting and turbine exhaust hoods are of welded sheet metal construction. Flanges utilizing dual (Naflex) seals are used at component connections. The purpose of the exhaust ducting is to conduct turbine exhaust gases to the thrust chamber exhaust manifold which encircles the thrust chamber approximately halfway between the throat and the nozzle exit. Exhaust gases pass through the heat exchanger and exhaust into the main thrust chamber through 180 triangular openings between the tubes of the thrust chamber.
HEAT EXCHANGER
The heat exchanger is a shell assembly, consisting of a duct, bellows, flanges, and coils. It is mounted in the turbine exhaust duet between the oxidizer turbopump and the thrust chamber. It heats and expands helium gas or converts liquid oxygen to gaseous oxygen for maintaining vehicle oxidizer tank pressurization. During engine operation, either liquid oxygen is tapped off the oxidizer high-pressure duct or helium is provided from vehicle stage and routed to the heat exchanger coils. As turbine exhaust gases pass over the heat exchanger coils, the liquid oxygen is converted to gaseous oxygen. The gaseous oxygen or helium is then routed to the vehicle oxidizer tank to maintain vehicle propellant tank pressurization.
OXIDIZER TURBINE BYPASS VALVE
The oxidizer turbine bypass valve is a normally open, spring-loaded, gate type. It is mounted in the oxidizer turbine bypass duct. The valve gate is equipped with a nozzle whose size is determined during engine calibration. The valve prevents an over-speed condition of the oxidizer turbopump and acts as a calibration device for the turbopump performance balance.
Control System
The control system includes a pneumatic system and a solid-state electrical sequence controller packaged with spark exciters for the gas generator and the thrust chamber spark plugs, plus interconnecting electrical cabling and pneumatic lines.
PNEUMATIC SYSTEM
The pneumatic system consists of a high pressure helium controlled gas storage tank, a regulator to reduce the pressure to a usable level, and electrical solenoid control valves to direct the central gas to the various pneumatically controlled valves.
ELECTRICAL SEQUENCE CONTROLLER
The electrical sequence controller is a completely self-contained, solid-state system, requiring only dc power and start and stop command signals. Pre-start status of all critical engine control functions is monitored in order to provide an "engine ready" signal. Upon obtaining "engine ready" and "start" signals, solenoid control valves are energized in a precisely timed sequence to bring the engine through ignition, transition, and into mainstage operation. After shutdown, the system automatically resets for a subsequent restart.
Start Tank Assembly System
This system is made up of an integral helium and hydrogen start tank, which contains the hydrogen and helium gases and the valves for starting and operating the engine.
HELIUM AND HYDROGEN TANKS
The spherical helium tank is positioned inside the hydrogen tank to minimize engine complexity. It holds 1000 cubic inches of helium. The larger spherical hydrogen gas tank has a capacity of 7257.6 cubic inches. Both tanks are filled from a ground source prior to launch.
Flight instrumentation System
The J-2 engine is provided with an instrumentation system which measures basic pressure, temperature, flow, speed and position parameters with the capability of transmitting these signals to a ground recording system or a telemetry system, or both. The instrumentation system is designed for use throughout the life of the J-2 engine, including the first static acceptance firing tests. The total instrumentation system consists of two independent primary and auxiliary packages. The primary package measures those parameters considered critical to all engine static firings and vehicle flights.
AUXILIARY PACKAGE
An auxiliary package is designed for use during early vehicle flights. It can be removed from the engine without interfering with the operation of the primary system. It contains sufficient flexibility to provide for deletion, substitution, or addition of parameters so that design changes requiring modifications will be minimized.
ENGINE OPERATION
Start Sequence
Start sequence is initiated by supplying energy to two spark plugs in the gas generator and to the augmented spark igniter to the thrust chamber for ignition of the propellants. Next, two solenoid valves are actuated: one for helium control, and one for ignition phase control. Helium is routed to hold the propellant bleed valves closed and to purge the thrust chamber LOX dome, the LOX pump intermediate seal, and the gas generator oxidizer passage. In addition, the main fuel and augmented spark igniter oxidizer valves are opened. An ignition flame is thus created in the center of the thrust chamber injector. A start tank discharge valve is then opened to initiate turbine spin. After a short interval of 0.64 seconds, this valve is closed and a mainstage control solenoid is actuated to: ( 1 ) turn off gas generator and thrust chamber helium purges; (2) open the gas generator control valve (hot gases from the gas generator now drive the pump turbines); (3) open the main oxidizer valve to the first position (14 degrees); (4) close the oxidizer turbine bypass valve (a portion of the gases for driving the oxidizer turbopump were bypassed during the ignition phase); (5) gradually bleed the pressure from the closing side of the oxidizer valve pneumatic actuator controlling the slow opening of this valve for smooth transition into mainstage. Energy in the spark plugs is cut off and the engine is operating at rated thrust. Within several seconds, the gaseous hydrogen tank will be recharged in those engines having a restart requirement. High pressure hydrogen gas is tapped from the fuel manifold located at the top of the thrust chamber to repressurize the tank.
Flight Mainstage Operation
During mainstage operation, engine thrust may be varied between 175,000 and 225,000 pounds by- actuating the propellant utilization valve to increase or decrease oxidizer flow. This is beneficial to some flight trajectories where overall mission performance is desired to make greater payloads possible.
Cutoff Sequence
When the engine cutoff signal is received by the electrical control package, it de-energizes the mainstage and ignition phase solenoid valves and energizes the helium control solenoid de-energizer timer. This, in turn, permits closing pressure to the fuel valve, oxidizer valve gas generator control valve, and augmented spark ignited valve. The oxidizer turbine bypass valve and propellant bleed valves open to complete the engine cutoff sequence.
Copyright 1997-2005 by John
Duncan |
