manu/mldr-zoomed-100char-total-queries-documents

chunk_id stringlengths 3 9	chunk stringlengths 1 100
9828_174	The turbine exhaust ducting and turbine exhaust hoods were of welded sheet metal construction.
9828_175	Flanges utilizing dual seals were used at component connections. The exhaust ducting conducted
9828_176	turbine exhaust gases to the thrust chamber exhaust manifold which encircled the combustion chamber
9828_177	approximately halfway between the throat and the nozzle exit. Exhaust gases passed through the heat
9828_178	exchanger and exhaust into the main combustion chamber through 180 triangular openings between the
9828_179	tubes of the combustion chamber.
9828_180	Heat exchanger
9828_181	The heat exchanger was a shell assembly, consisting of a duct, bellows, flanges, and coils. It was
9828_182	mounted in the turbine exhaust duct between the oxidizer turbine discharge manifold and the thrust
9828_183	chamber. It heated and expanded helium gas for use in the third stage or converted LOX to gaseous
9828_184	oxygen for the second stage for maintaining vehicle oxidizer tank pressurization. During engine
9828_185	operation, either LOX was tapped off the oxidizer high-pressure duct or helium was provided from
9828_186	the vehicle stage and routed to the heat exchanger coils.
9828_187	Start tank assembly system
9828_188	This system was made up of an integral helium and hydrogen start tank, which contained the hydrogen
9828_189	and helium gases for starting and operating the engine. The gaseous hydrogen imparted initial spin
9828_190	to the turbines and pumps prior to gas generator combustion, and the helium was used in the control
9828_191	system to sequence the engine valves. The spherical helium tank was positioned inside the hydrogen
9828_192	tank to minimize engine complexity. It held of helium. The larger spherical hydrogen gas tank had
9828_193	a capacity of . Both tanks were filled from a ground source prior to launch and the gaseous
9828_194	hydrogen tank was refilled during engine operation from the thrust chamber fuel inlet manifold for
9828_195	subsequent restart in third stage application.
9828_196	Control system
9828_197	The control system included a pneumatic system and a solid-state electrical sequence controller
9828_198	packaged with spark exciters for the gas generator and the thrust chamber spark plugs, plus
9828_199	interconnecting electrical cabling and pneumatic lines, in addition to the flight instrumentation
9828_200	system. The pneumatic system consisted of a high-pressure helium gas storage tank, a regulator to
9828_201	reduce the pressure to a usable level, and electrical solenoid control valves to direct the central
9828_202	gas to the various pneumatically controlled valves. The electrical sequence controller was a
9828_203	completely self-contained, solid-state system, requiring only DC power and start and stop command
9828_204	signals. Pre-start status of all critical engine control functions was monitored in order to
9828_205	provide an "engine ready" signal. Upon obtaining "engine ready" and "start" signals, solenoid
9828_206	control valves were energized in a precisely timed sequence to bring the engine through ignition,
9828_207	transition, and into main-stage operation. After shutdown, the system automatically reset for a
9828_208	subsequent restart.
9828_209	Flight instrumentation system
9828_210	The flight instrumentation system is composed of a primary instrumentation package and an auxiliary
9828_211	package. The primary package instrumentation measures those parameters critical to all engine
9828_212	static firings and subsequent vehicle launches. These include some 70 parameters such as pressures,
9828_213	temperatures, flows, speeds, and valve positions for the engine components, with the capability of
9828_214	transmitting signals to a ground recording system or a telemetry system, or both. The
9828_215	instrumentation system is designed for use throughout the life of the engine, from the first static
9828_216	acceptance firing to its ultimate vehicle flight. The auxiliary package is designed for use during
9828_217	early vehicle flights. It may be deleted from the basic engine instrumentation system after the
9828_218	propulsion system has established its reliability during research and development vehicle flights.
9828_219	It contains sufficient flexibility to provide for deletion, substitution, or addition of parameters
9828_220	deemed necessary as a result of additional testing. Eventual deletion of the auxiliary package will
9828_221	not interfere with the measurement capability of the primary package.
9828_222	Engine operation
9828_223	Start sequence
9828_224	Start sequence was initiated by supplying energy to two spark plugs in the gas generator and two in
9828_225	the augmented spark igniter for ignition of the propellants. Next, two solenoid valves were
9828_226	actuated; one for helium control, and one for ignition phase control. Helium was routed to hold the
9828_227	propellant bleed valves closed and to purge the thrust chamber LOX dome, the LOX pump intermediate
9828_228	seal, and the gas generator oxidizer passage. In addition, the main fuel and ASI oxidizer valves
9828_229	were opened, creating an ignition flame in the ASI chamber that passed through the center of the
9828_230	thrust chamber injector.
9828_231	After a delay of 1, 3, or 8 seconds, during which time fuel was circulated through the thrust
9828_232	chamber to condition the engine for start, the start tank discharge valve was opened to initiate
9828_233	turbine spin. The length of the fuel lead was dependent upon the length of the Saturn V first stage
9828_234	boost phase. When the engine was used in the S-II stage, a 1-second fuel lead was necessary. The
9828_235	S-IVB, on the other hand, utilized a 1-second fuel lead for its initial start and an 8-second fuel
9828_236	lead for its restart.
9828_237	After an interval of 0.450 seconds, the start tank discharge valve was closed and a mainstage
9828_238	control solenoid was actuated to:
9828_239	Turn off gas generator and thrust chamber helium purges
9828_240	Open the gas generator control valve (hot gases from the gas generator now drive the pump turbines)
9828_241	Open the main oxidizer valve to the first position (14 degrees) allowing LOX to flow to the LOX
9828_242	dome to burn with the fuel that has been circulating through the injector
9828_243	Close the oxidizer turbine bypass valve (a portion of the gases for driving the oxidizer turbopump
9828_244	were bypassed during the ignition phase)
9828_245	Gradually bleed the pressure from the closing side of the oxidizer valve pneumatic actuator
9828_246	controlling the slow opening of this valve for smooth transition into mainstage.
9828_247	Energy in the spark plugs was cut off and the engine was operating at rated thrust. During the
9828_248	initial phase of engine operation, the gaseous hydrogen start tank would be recharged in those
9828_249	engines having a restart requirement. The hydrogen tank was repressurized by tapping off a
9828_250	controlled mixture of LH2 from the thrust chamber fuel inlet manifold and warmer hydrogen from the
9828_251	thrust chamber fuel injection manifold just before entering the injector.
9828_252	Flight mainstage operation
9828_253	During mainstage operation, engine thrust could be varied between by actuating the propellant
9828_254	utilization valve to increase or decrease oxidizer flow. This was beneficial to flight trajectories
9828_255	and for overall mission performance to make greater payloads possible.
9828_256	Cutoff sequence
9828_257	When the engine cutoff signal was received by the electrical control package, it de-energized the
9828_258	main-stage and ignition phase solenoid valves and energized the helium control solenoid
9828_259	de-energizer timer. This, in turn, permitted closing pressure to the main fuel, main oxidizer, gas
9828_260	generator control, and augmented spark igniter valves. The oxidizer turbine bypass valve and
9828_261	propellant bleed valves opened and the gas generator and LOX dome purges were initiated.
9828_262	Engine restart
9828_263	To provide third stage restart capability for the Saturn V, the J-2 gaseous hydrogen start tank was
9828_264	refilled in 60 seconds during the previous firing after the engine had reached steady-state
9828_265	operation (refill of the gaseous helium tank was not required because the original ground-fill
9828_266	supply was sufficient for three starts). Prior to engine restart, the stage ullage rockets were
9828_267	fired to settle the propellants in the stage propellant tanks, ensuring a liquid head to the
9828_268	turbopump inlets. In addition, the engine propellant bleed valves were opened, the stage
9828_269	recirculation valve was opened, the stage prevalve was closed, and a LOX and LH2 circulation was
9828_270	effected through the engine bleed system for five minutes to condition the engine to the proper
9828_271	temperature to ensure proper engine operation. Engine restart was initiated after the "engine
9828_272	ready" signal was received from the stage. This was similar to the initial "engine ready". The hold
9828_273	time between cutoff and restart was from a minimum of 1.5 hours to a maximum of 6 hours, depending

9828_174

The turbine exhaust ducting and turbine exhaust hoods were of welded sheet metal construction.

9828_175

Flanges utilizing dual seals were used at component connections. The exhaust ducting conducted

9828_176

turbine exhaust gases to the thrust chamber exhaust manifold which encircled the combustion chamber

9828_177

approximately halfway between the throat and the nozzle exit. Exhaust gases passed through the heat

9828_178

exchanger and exhaust into the main combustion chamber through 180 triangular openings between the

9828_179

tubes of the combustion chamber.

9828_180

Heat exchanger

9828_181

The heat exchanger was a shell assembly, consisting of a duct, bellows, flanges, and coils. It was

9828_182

mounted in the turbine exhaust duct between the oxidizer turbine discharge manifold and the thrust

9828_183

chamber. It heated and expanded helium gas for use in the third stage or converted LOX to gaseous

9828_184

oxygen for the second stage for maintaining vehicle oxidizer tank pressurization. During engine

9828_185

operation, either LOX was tapped off the oxidizer high-pressure duct or helium was provided from

9828_186

the vehicle stage and routed to the heat exchanger coils.

9828_187

Start tank assembly system

9828_188

This system was made up of an integral helium and hydrogen start tank, which contained the hydrogen

9828_189

and helium gases for starting and operating the engine. The gaseous hydrogen imparted initial spin

9828_190

to the turbines and pumps prior to gas generator combustion, and the helium was used in the control

9828_191

system to sequence the engine valves. The spherical helium tank was positioned inside the hydrogen

9828_192

tank to minimize engine complexity. It held of helium. The larger spherical hydrogen gas tank had

9828_193

a capacity of . Both tanks were filled from a ground source prior to launch and the gaseous

9828_194

hydrogen tank was refilled during engine operation from the thrust chamber fuel inlet manifold for

9828_195

subsequent restart in third stage application.

9828_196

Control system

9828_197

The control system included a pneumatic system and a solid-state electrical sequence controller

9828_198

packaged with spark exciters for the gas generator and the thrust chamber spark plugs, plus

9828_199

interconnecting electrical cabling and pneumatic lines, in addition to the flight instrumentation

9828_200

system. The pneumatic system consisted of a high-pressure helium gas storage tank, a regulator to

9828_201

reduce the pressure to a usable level, and electrical solenoid control valves to direct the central

9828_202

gas to the various pneumatically controlled valves. The electrical sequence controller was a

9828_203

completely self-contained, solid-state system, requiring only DC power and start and stop command

9828_204

signals. Pre-start status of all critical engine control functions was monitored in order to

9828_205

provide an "engine ready" signal. Upon obtaining "engine ready" and "start" signals, solenoid

9828_206

control valves were energized in a precisely timed sequence to bring the engine through ignition,

9828_207

transition, and into main-stage operation. After shutdown, the system automatically reset for a

9828_208

subsequent restart.

9828_209

Flight instrumentation system

9828_210

The flight instrumentation system is composed of a primary instrumentation package and an auxiliary

9828_211

package. The primary package instrumentation measures those parameters critical to all engine

9828_212

static firings and subsequent vehicle launches. These include some 70 parameters such as pressures,

9828_213

temperatures, flows, speeds, and valve positions for the engine components, with the capability of

9828_214

transmitting signals to a ground recording system or a telemetry system, or both. The

9828_215

instrumentation system is designed for use throughout the life of the engine, from the first static

9828_216

acceptance firing to its ultimate vehicle flight. The auxiliary package is designed for use during

9828_217

early vehicle flights. It may be deleted from the basic engine instrumentation system after the

9828_218

propulsion system has established its reliability during research and development vehicle flights.

9828_219

It contains sufficient flexibility to provide for deletion, substitution, or addition of parameters

9828_220

deemed necessary as a result of additional testing. Eventual deletion of the auxiliary package will

9828_221

not interfere with the measurement capability of the primary package.

9828_222

Engine operation

9828_223

Start sequence

9828_224

Start sequence was initiated by supplying energy to two spark plugs in the gas generator and two in

9828_225

the augmented spark igniter for ignition of the propellants. Next, two solenoid valves were

9828_226

actuated; one for helium control, and one for ignition phase control. Helium was routed to hold the

9828_227

propellant bleed valves closed and to purge the thrust chamber LOX dome, the LOX pump intermediate

9828_228

seal, and the gas generator oxidizer passage. In addition, the main fuel and ASI oxidizer valves

9828_229

were opened, creating an ignition flame in the ASI chamber that passed through the center of the

9828_230

thrust chamber injector.

9828_231

After a delay of 1, 3, or 8 seconds, during which time fuel was circulated through the thrust

9828_232

chamber to condition the engine for start, the start tank discharge valve was opened to initiate

9828_233

turbine spin. The length of the fuel lead was dependent upon the length of the Saturn V first stage

9828_234

boost phase. When the engine was used in the S-II stage, a 1-second fuel lead was necessary. The

9828_235

S-IVB, on the other hand, utilized a 1-second fuel lead for its initial start and an 8-second fuel

9828_236

lead for its restart.

9828_237

After an interval of 0.450 seconds, the start tank discharge valve was closed and a mainstage

9828_238

control solenoid was actuated to:

9828_239

Turn off gas generator and thrust chamber helium purges

9828_240

Open the gas generator control valve (hot gases from the gas generator now drive the pump turbines)

9828_241

Open the main oxidizer valve to the first position (14 degrees) allowing LOX to flow to the LOX

9828_242

dome to burn with the fuel that has been circulating through the injector

9828_243

Close the oxidizer turbine bypass valve (a portion of the gases for driving the oxidizer turbopump

9828_244

were bypassed during the ignition phase)

9828_245

Gradually bleed the pressure from the closing side of the oxidizer valve pneumatic actuator

9828_246

controlling the slow opening of this valve for smooth transition into mainstage.

9828_247

Energy in the spark plugs was cut off and the engine was operating at rated thrust. During the

9828_248

initial phase of engine operation, the gaseous hydrogen start tank would be recharged in those

9828_249

engines having a restart requirement. The hydrogen tank was repressurized by tapping off a

9828_250

controlled mixture of LH2 from the thrust chamber fuel inlet manifold and warmer hydrogen from the

9828_251

thrust chamber fuel injection manifold just before entering the injector.

9828_252

Flight mainstage operation

9828_253

During mainstage operation, engine thrust could be varied between by actuating the propellant

9828_254

utilization valve to increase or decrease oxidizer flow. This was beneficial to flight trajectories

9828_255

and for overall mission performance to make greater payloads possible.

9828_256

Cutoff sequence

9828_257

When the engine cutoff signal was received by the electrical control package, it de-energized the

9828_258

main-stage and ignition phase solenoid valves and energized the helium control solenoid

9828_259

de-energizer timer. This, in turn, permitted closing pressure to the main fuel, main oxidizer, gas

9828_260

generator control, and augmented spark igniter valves. The oxidizer turbine bypass valve and

9828_261

propellant bleed valves opened and the gas generator and LOX dome purges were initiated.

9828_262

Engine restart

9828_263

To provide third stage restart capability for the Saturn V, the J-2 gaseous hydrogen start tank was

9828_264

refilled in 60 seconds during the previous firing after the engine had reached steady-state

9828_265

operation (refill of the gaseous helium tank was not required because the original ground-fill

9828_266

supply was sufficient for three starts). Prior to engine restart, the stage ullage rockets were

9828_267

fired to settle the propellants in the stage propellant tanks, ensuring a liquid head to the

9828_268

turbopump inlets. In addition, the engine propellant bleed valves were opened, the stage

9828_269

recirculation valve was opened, the stage prevalve was closed, and a LOX and LH2 circulation was

9828_270

effected through the engine bleed system for five minutes to condition the engine to the proper

9828_271

temperature to ensure proper engine operation. Engine restart was initiated after the "engine

9828_272

ready" signal was received from the stage. This was similar to the initial "engine ready". The hold

9828_273

time between cutoff and restart was from a minimum of 1.5 hours to a maximum of 6 hours, depending