Category Apollo Saturn V News Reference

Filght Control System

The flight control system provides stage thrust vector steering and attitude control. Steering is achieved by gimbaling the J-2 engine during pow-

Filght Control System

5-8

 

SATURN V NEWS

ered flight. Hydraulic actuator assemblies provide J-2 engine deflection rates proportional to steering signal corrections supplied by the IU.

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Stage roll attitude during powered flight is con­trolled by firing the APS attitude control engines.

MOBILE LAUNCHER

The mobile launcher is a movable launch platform with an integral umbilical tower. The launcher base

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K-107-66PC-91

Arrival to Launch Pad—The facilities vehicle arrives at Launch Complex 39A.

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is a two-story steel structure covering more than half an acre. The 380-foot tower, which supports the electrical servicing and fluid lines for the ve­hicle, is a steel structure mounted on the base. The base and tower weigh 10.5 million pounds and stand 445 feet above ground level.

Among major considerations in design of the mobile launcher were crew safety and escape provisions

and protection of the platform and its equipment from blast and sonic damage.

Personnel may be evacuated from upper work levels of the umbilical tower by a high speed elevator, descending at 600 feet per minute. After leaving the elevator, they can drop through a flexible metal chute into a blast and heatproof room inside the base of the pad hardstand.

The mobile launcher provides physical support and is a major facility for checkout of the space vehicle from assembly at the VAB until liftoff at the launch

site.

The top level of the launcher base houses digital acquisition units, computer systems, controls for actuation of service arms, communications equip­ment, water deluge panels, and other control units. Included in the lower level are hydraulic charging units, environmental control systems, electrical measuring equipment, and a terminal room for in­strumentation and communications interface. Mounted on the top deck of the base are four vehicle holddown and support arms and three tail service masts.

The umbilical tower is an open steel structure pro­viding support for nine umbilical service arms, 18 work and access platforms, and, for propellant, pneumatic, electrical, water, communications, and other service lines required to sustain the vehicle. A 250-ton capacity hammerhead crane is mounted atop the umbilical tower.

The launcher restrains the vehicle for approximately 5 seconds after ignition to allow thrust buildup and verification of full thrust from all engines. The design “up-load” during the holddown period is 3 million pounds. If one or more of the engines fail to develop full thrust, the vehicle is not released, and all engines automatically are shut down.

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Night Shot—A 365-foot-tall Saturn V facilities vehicle Is shown in place at Launch Pad 39A.

SATURN V NEWS REFERENCE

THE SATURN V

INTRODUCTION

When the United States made the decision in 1961 to undertake a manned lunar landing effort as the focal point of a broad new space exploration pro­gram, there was no rocket in the country even approaching the needed capability. There was a sort of “test bed” in the making, a multi-engine vehicle now known as Saturn I. It had never flown. And it was much too small to offer any real hope of sending a trio to the moon, except possibly through as many as a half dozen separate launchings from earth and the perfection of rendezvous and docking techniques, which had never been tried.

That was the situation that brought about the an­nouncement on Jan. 10, 1962, that the National Aeronautics and Space Administration would de­velop a new rocket, much larger than any previously attempted. It would be based on the F-l rocket en­gine, the development of which had been underway since 1958, and the hydrogen-fueled J-2 engine, upon which work had begun in 1960.

The Saturn V, then, is the first large vehicle in the U. S. space program to be conceived and de­veloped for a specific purpose. The lunar landing task dictated the make-up of the vehicle, but it was not developed solely for that mission. As President Kennedy pointed out when he issued his space chal­lenge to the Congress on May 25, 1961, the overall objective is for “this Nation to take a clearly lead­ing role in space achievement which in many ways may hold the key to our future on earth.” He said of the lunar landing project: “No single space pro­ject in this period will be more exciting, or more impressive to mankind, or more important for the long-range exploration of space: and none will be so difficult or expensive to accomplish…”

The Saturn V program is the biggest rocket effort undertaken in this country. Its total cost, including the production of 15 vehicles between now and early 1970, will be above $7 billion.

NASA formally assigned the task of developing the Saturn V to the Marshall Space Flight Center on Jan. 25, 1962. Launch responsibility was committed to the Kennedy Space Center. (The Manned Space­craft Center, the third center in manned space flight, is responsible for spacecraft development, crew training, and inflight control.)

DESCRIPTION

Marshall Center rocket designers conceived the Saturn V in 1961 and early 1962. They decided that
a three-stage vehicle would best serve the immedi­ate needs for a lunar landing mission and would serve well as a general purpose space exploration vehicle.

One of the more important decisions made early in the program called for the fullest possible use of components and techniques proven in the Saturn I program. As a result, the Saturn V third stage (S-IVB) was patterned after the Saturn I second stage (S-IV). And the Saturn V instrument unit is an outgrowth of the one used on Saturn I. In these areas, maximum use of designs and facilities already avail­able was incorporated to save time and costs.

Many other components were necessary, including altogether new first and second stages (S-IC and S-II). The F-l and J-2 engines were already under development, although much work remained to be done. The guidance system was to be an improve­ment on that of the Saturn I.

Saturn V, including the Apollo spacecraft, is 364 feet tall. Fully loaded, the vehicle will weigh some

6.1 million pounds.

The 300,000-pound first stage is 33 feet in diameter and 138 feet long. It is powered by five F-l engines generating 7.5 million pounds thrust. The booster will burn 203,000 gallons of RP-1 (refined kerosene) and 331,000 gallons of liquid oxygen (LOX) in 2.5 minutes.

Saturn V’s second stage is powered by five J-2 engines that generate a total thrust of a million pounds. The 33-foot diameter stage weighs 95,000 pounds empty and more than a million pounds loaded. It burns some 260,000 gallons of liquid hydrogen and

83.0 gallons of liquid oxygen during a typical 6- minute flight.

Third stage of the vehicle is 21 feet and 8 inches in diameter and 58 feet and 7 inches long. An inter­stage adapter connects the larger diameter second stage to the smaller upper stage. Empty weight of the stage is 34,000 pounds and the fueled weight is

262.0 pounds. A single J-2 engine developing up to 225,000 pounds of thrust powers the stage. Typi­cal burn time is 2.75 minutes for the first burn and

5.2 minutes to a translunar injection.

The vehicle instrument unit sits atop the third stage. The unit, which weighs some 4,500 pounds, contains the electronic gear that controls engine ig­nition and cutoff, steering, and all other commands necessary for the Saturn V mission. Diameter of the instrument unit is 21 feet and 8 inches, and height is 3 feet.

Directly above the instrument unit in the Apollo

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configuration is the Apollo spacecraft. It consists of the lunar module, the service module, the com­mand module, and the launch escape system. Total height of the package is about 80 feet.

ENGINE INTERFACE PANEL

The engine interface panel, mounted above the turbopump LOX and fuel inlets, provides the ve­hicle connect location for electrical connectors be­tween the engine and the vehicle. It also provides the attachment point for the vehicle flexible heat – resistant curtain. The panel is fabricated from heat – resistant stainless-steel casting made in three sec­tions and assembled by rivets and bolts.

ELECTRICAL SYSTEM

The electrical system consists of flexible armored wiring harnesses for actuation of engine controls and the flight instrumentation harnesses.

HYDRAULIC CONTROL SYSTEM

The hydraulic control system operates the engine propellant valves during the start and cutoff se­quences. It consists of a hypergol manifold, a check­out valve, an engine control valve, and the related tubing and fittings.

Hypergol Manifold

The hypergol manifold directs hypergolic fluid to the separate igniter fuel system in the thrust cham­ber injector. It consists of a hypergol container, an ignition monitor valve, a position switch, and an igniter fuel valve. The hypergol container, position switch, and igniter fuel valve are internal parts of the hypergol manifold.

A spring-loaded, cam-lock mechanism incorporated in the hypergol manifold prevents actuation of the
ignition monitor valve until after the upstream hypergol cartridge diaphragm bursts. The same mechanism actuates a position switch that indicates when the hypergol cartridge is installed. The igniter fuel valve is a spring-loaded, cracking check valve that opens and allows fuel to flow into the hypergol container. The hypergol cartridge diaphragms are ruptured by the resultant pressure surge when the igniter fuel valve opens.

Ignition Monitor Valve

The ignition monitor valve is a pressure-actuated, three-way valve mounted on the hypergol mani­fold. It controls the opening of the fuel valves and permits them to open only after satisfactory com­bustion has been achieved in the thrust chamber.

When the hypergol cartridge is installed in the hypergol manifold, a cam-lock mechanism prevents the ignition monitor valve poppet from moving from the closed position. The ignition monitor valve has six ports: a control port, an inlet port, two outlet ports, a return port, and an atmospheric reference port. The control port receives pressure from the thrust chamber fuel manifold. The inlet port re­ceives hydraulic fuel pressure for opening the fuel valves. When the ignition monitor valve poppet is in the deactuated position, hydraulic fuel from the inlet port is stopped at the poppet seat. When the hypergol cartridge diaphragm bursts, the spring- loaded cam-lock retracts to permit the ignition moni­tor valve poppet unrestricted motion. When thrust chamber pressure (directed to the control port from the thrust chamber fuel manifold I increases, the ignition monitor valve poppet moves to the open (actuated) position and hydraulic fuel is directed through the outlet ports to the fuel valves.

Gas Generator and Exhaust System

This system consists of the gas generator, gas gen­erator control valve, turbine exhaust system and exhaust manifold, heat exchanger, and oxidizer turbine bypass valve.

GAS GENERATOR

The gas generator is welded to the fuel pump tur­bine manifold, making it an integral part of the fuel turbopump assembly. It produces hot gases to drive the fuel and oxidizer turbines and consists of a combustor containing two spark plugs, a control valve containing fuel and oxidizer 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 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 fuel and oxidizer 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 exhaust ducting con­ducts 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 duct between the oxidizer turbine discharge manifold and the thrust chamber. It heats and expands helium gas for use in the third stage or converts liquid oxygen to gaseous oxygen for the second stage 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 the ve­hicle stage and routed to the heat exchanger coils.

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 w’ith a nozzle, the size of which is deter­mined during engine calibration. The valve in its open position depresses the speed of the oxygen pump during start, and in its closed position 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 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.

FUEL FEED SYSTEM

Ten fuel suction lines (two per engine) supply fuel from the fuel tank to the five F-l engines. The suc­tion line outlets attach directly to the F-l engine fuel pump inlets.

Each suction line has a pneumatically controlled fuel prevalve which normally remains open. This

image37Подпись:Подпись:image38"FUEL-CONDITIONING (BUBBLING) SYSTEM

The fuel-conditioning system bubbles gaseous ni­trogen through the fuel feed lines and fuel tank to prevent fuel temperature stratification prior to launch. A wire mesh filter in the nitrogen supply line prevents discharge of contaminants into the conditioning system.

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Fuel Conditioning

A check valve in the outlet of each fuel-conditioning line prevents fuel from entering the nitrogen lines when the fuel-conditioning system is not operating.

An orifice located near each fuel-conditioning check valve provides the proper nitrogen flow into each fuel duct.

PROPELLANT SYSTEM

The propellant system is composed of seven sub­systems : purge, fill and replenish, venting, pres­surization, propellant feed, recirculation, and pro­pellant management..

Purge Subsystem

The purge subsystem uses helium gas to clear the propellant tanks of contaminants before they are loaded. The important contaminants art* oxygen in the liquid hydrogen tank (liquid hydrogen will freeze oxygen which is impact-sensitive) and moisture in the liquid oxygen tank.

The tanks are purged with helium gas from ground storage tanks. The tanks are alternately pressur­ized and vented to dilute the concentration of con­taminants. The operation is repeated until samples of the helium gas emptied from the tanks show that contaminants have been removed or reduced to a safe level.

Fill and Replenish Subsystem

Filling of the propellant tanks on the second stage is a complex and precise task because of the nature of thd cryogenic liquids and the construction of the stage.

Because the metal of the stage is at normal outside temperature, it must be chilled gradually before pumping the ultra-cold propellants into the tanks. The filling operation thus starts with the introduc­tion of cold gas into the tanks, lines, valves, and other components that will come into contact with the cryogenic fluids. The cold gas is circulated until

4.7

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image86"S-19

Channel Installed—Feed line from IH2 tank to one of the five engines is installed.

the metal has become chilled enough to begin pump­ing in the propellants. The filling and replenishing subsystem operation consists of five phases:

Chilldown – Propellants are first pumped into the tank at the rate of 500 gallons per minute for LQX and 1,000 gallons per minute for LHa. Despite the preliminary chilling by cold gas, the tanks are still so much warmer than the propellants that much of the latter boils off (converts to gaseous form) when it first goes into the tank. Filling con­tinues at this rate until enough of the propellants remain liquid so that the tanks are full to the five per cent level.

Fast Fill —As soon as tank sensors report that the liquid has reached the five per cent level, the fill­ing rate is increased to 5,000 gallons per minute for LOX and 10,000 gallons per minute for LH2. This rate continues until the liquid level in the tank reaches the 98 per cent level.

Slow-Fill—Propellant tanks are filled at the rate of 1,000 gallons per minute for both LOX and LH2 until the 100 per cent level is reached.

Replenishment—Because filling begins many hours before a scheduled liftoff and the cryogenic liquids are constantly boiling off, filling continues almost up to liftoff (160 seconds before liftoff for LOX and 70 seconds before liftoff for LH2). Tanks

are filled at the rate of up to 200 gallons per min­ute for LOX and up to 500 gallons per minute for LH,, depending on signals from sensors in the tanks on the liquid level.

101 Per Cent Shutdown—A sensor in each tank will send a signal to indicate that the 101 per cent level (over the proper fill level) has been reached; this signal causes immediate shutdown of filling operations.

Filling is accomplished through separate connec­tions, lines, and valves. The ground part of the con­nections is covered by special shrouds in which he­lium is circulated during filling operations. This provides an inert atmosphere around the coupling between the ground line and the tanks.

The coupling of the fill line and the tanks is engaged manually at the start of filling operations; it is nor­mally disengaged remotely by applying pneumatic pressure to the coupling lock and actuating a push – off mechanism. A backup method involves a remotely attached lanyard in which the vertical rise of the vehicle will unlock the coupling. The fill valves are designed so that loss of helium pressure or electrical power will automatically close them.

Liquid oxygen is the first propellant to be loaded onto the stage. It is pumped from ground storage tanks. Liquid hydrogen is transferred to the stage by pressurizing the ground storage tanks with gaseous hydrogen. The liquid hydrogen tank is chilled before the liquid oxygen is loaded to avoid structural stresses.

After filling is completed, the fill valves and the liquid oxygen debris valves in the coupling are closed, but the liquid hydrogen debris valve is left open. The liquid oxygen fill line is then drained and purged with helium. The liquid hydrogen line is purged up to the coupling. When a certain signal is received (first stage thrust-commit), the liquid hydrogen debris valve is closed and the coupling is separated from the stage.

The tanks can be drained by pressurizing them, opening the valves, and reversing the filling opera­tion.

IBM FACILITIES

Three IBM-owned buildings at Huntsville comprise the Space Systems Center where component test­ing, fabrication, assembly, and systems checkout of the instrument unit are completed. Assembly and the majority of the testing activity take place in a 130,000-square-foot building located in Hunts­ville’s Research Park.

As units are received, they are inspected and then moved to one of the testing laboratories where they are subjected to detailed quality and reliability testing. From component testing, the parts move

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IBM-DR-24

Ш Assembly and Test—All instrument unit assembly work and the majority of testing are done in this IBM-owned building in Huntsville’s Research Park. The rear of the building is the high – bay area where assembly operations take place.

Подпись:Following assembly operations, the IU is moved to one of two systems checkout stands—one for uprated Saturn I vehicles, the other for Saturn V.

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IBM-DR-25

Automatic Checkout-IBM technicians monitor systems checkout tests as another technician optically adjusts the inertial guidance platform, prior to a simulated mission.

A complete systems checkout is performed auto­matically. Hooked by underground cables, two digital checkout computer systems examine the IU. Each of the IU’s six subsystems is tested before the IU is tested as an integrated unit. With indepen­dent computers, systems tests for two instrument units can be conducted simultaneously.

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IBMDR-26

Simulation Laboratory Saturn V flight guidance and navigation programs as well as launch computer programs are tested in IBM’s Engineering Building at Huntsville. Here a technician checks a computer readout of a simulated mission.

FIRST STAGE FLIGHT

The first stage is loaded with RP-1 fuel and LOX at approximately 12 and 4 hours respectively, be­fore launch. With all systems in a ready condition, the stage is ignited by sending a start signal to the five F-l rocket engines. The engine main LOX valves open first allowing LOX to begin to enter the main thrust chamber. Next the engines’ gas generators and turbopumps are started. Each en­gine’s turbopump assembly will develop approx­imately 60,000 horsepower. Combustion is initiated by injecting a hypergolic solution into the engine’s main thrust chamber to react with the LOX already present. The main fuel valves then open, and fuel enters the combustion chamber to sustain the re­action previously initiated by the LOX and hyper­golic solution. Engine thrust then rapidly builds up to full level. The five engines are started in a 1-2-2 sequence, the center engine first and opposing out­board pairs at 300-millisecond stagger times. The stage is held down while the engines build up full thrust. After full thrust is reached and all engines and stage systems are functioning properly, the stage is released. This is accomplished by a “soft release” mechanism. First, the restraining hold­down arms are released. Immediately thereafter the vehicle begins to ascend but with a restraining force caused by tapered metal pins being pulled through holes. This “soft release” lasts for about 500 milliseconds.

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The vehicle rises vertically to an altitude of approx­imately 430 feet to clear the launch umbilical tower and then begins a pitch and roll maneuver to attain the correct flight azimuth. As the vehicle continues its flight, its path is controlled by gimbaling the outboard F-l engines consistent with a prepro­grammed flight path and commanded by the instru­ment unit.

At approximately 09 seconds into the flight, the vehicle experiences a condition of maximum dy­namic pressure. At this time, the restraining drag force is approximately equal to 400,000 pounds.

At 135.5 seconds into the flight most of the LOX and fuel will be consumed, and a signal is sent from the instrument unit to shut down the center engine. The outboard engines continue to burn until either LOX or fuel depletion is sensed. LOX depletion is signaled w’hen a “dry” indication is received from at least two of the four LOX cutoff sensors; one sensor is located near the top of each outboard LOX suction duct. Fuel depletion is signaled by a “dry” indication from a redundant fuel cutoff sensor bolted directly to the fuel tank lower bulkhead. The LOX depletion cutoff is the main cutoff system with fuel cutoff as the backup.

Six hundred milliseconds after the outboard engines receive a cutoff signal, a signal is given to fire the first stage retrorockets. Eight retroroekets are pro­vided and each produces an average effective thrust of 88,600 pounds for 0.666 seconds. The first stage separates from the second stage at an altitude of about 205,000 feet. It then ascends to a peak altitude near 366,000 feet before beginning its descent. While falling, the stage assumes a semistable en­gines down position and impacts into the Atlantic Ocean at approximately 350 miles down range of Cape Kennedy.

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F-l ENGINE FACT SHEET

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Подпись: NOTE: E-l engine will be uprated to 1,522,000 ib. thrust for Vehicle 504 and all subsequent operational vehicles.

LENGTH

WIDTH

THRUST (sea level)

SPECIFIC IMPULSE (minimum)

RATED RUN DURATION FLOWRATE: Oxidizer Fuel

MIXTURE RATIO CHAMBER PRESSURE WEIGHT FLIGHT CONFIGURATION EXPANSION AREA RATIO

COMBUSTION TEMPERATURE: Thrust Chamber

Gas Generator

MAXIMUM NOZZLE EXIT DIAMETER

19 ft.

12 (t. 4 in.

1.500,0 lb.

260 sec.

150 sec.

3,945 lb. sec. (24,811 gpm) 1,738 lb. sec. (15,471 gpm) 2,27:1 oxidizer to fuel 965 psia

18,500 lb. maximum

16:1 with nozzle extension

10:1 without nozzle extension

5,970°F

1,465°F

11 ft. 7 in,

^ SATURN V NEWS REFERENCE

HYDRAULIC SYSTEM

The hydraulic system performs engine positioning upon command from the IU. Major components are a J-2 engine-driven hydraulic pump, two hydraulic actuator assemblies, and an accumulator-reservoir assembly.

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J-2 Engine Hydraulic System Components

The electrically driven auxiliary hydraulic pump is started before vehicle liftoff to pressurize the hy­draulic system. Electric power for the pump is provided by a ground source. At liftoff, the pump is switched to stage battery power. Pressurization of the hydraulic system restrains the J-2 engine in a null position with relation to the third stage eenter-

REFERENCE

line, preventing pendulum-like shifting from forces encountered during liftoff and boost. During power­ed flight, the J-2 engine may be gimbaled up to 7° in a square pattern by the hydraulic system upon command from the IU.

Engine-Driven Hydraulic Pump

The engine-driven hydraulic pump is a variable dis­placement type pump capable of delivering hy­draulic fluid under continuous system pressure and varying volume as required for operation of the hy­draulic actuator assemblies. The pump is driven directly from the engine oxidizer turbopump. A thermal isolator in the system controls hydraulic – fluid temperature to ensure proper operation.

Auxiliary Hydraulic Pump

The auxiliary hydraulic pump is an electrically driven variable displacement pump which supplies a constant minimum supply of hydraulic fluid to the hydraulic system at all times. The pump is also used to perform preflight engine gimbaling check­outs, hydraulically lock the engine in the null posi­tion during boost phase, maintain system hydraulic – fluid at operating temperatures during other than the powered phase, and augment the engine-driven hydraulic pump during powered flight. It also pro­vides an emergency backup supply of fluid to the system.

Hydraulic Actuator Assemblies

Two hydraulic actuator assemblies are attached directly to the J-2 engine and the thrust structure and receive IU command signals to gimba! the en­gine. The actuator assemblies are identical and interchangeable.

Accumulator-Reservoir Assembly

The accumulator-reservoir assembly is an integral unit mounted on the thrust structure. The reservoir section is the storage area for hydraulic fluid; the accumulator section supplies peak system fluid re­quirements and dampens high-pressure surges with­in the system.