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

chunk_id stringlengths 3 9	chunk stringlengths 1 100
9827_374	to give everything they had to this cause. However, when they failed to live the Law of
9827_375	Consecration, which was a promise to voluntarily give all their property to the community for equal
9827_376	distribution, Joseph Smith received a revelation that God would build Zion later, when the people
9827_377	were ready. Both the Community of Christ and the LDS Church have come to understand Zion as
9827_378	primarily a community way of life that is centered around Jesus Christ, but many people still
9827_379	believe that, when they are ready, they will be able to build the literal city of Zion as well.
9827_380	See also
9827_381	Cunning Folk Traditions and the Latter Day Saint Movement
9827_382	History of The Church of Jesus Christ of Latter-day Saints Mormonism in the 19th century
9827_383	Death in 19th-century Mormonism Mormonism in the 20th century Mormonism in the 21st century
9827_384	Origin of Latter Day Saint polygamy Latter Day Saint Historians
9827_385	Criticism of The Church of Jesus Christ of Latter-day Saints Mormonism and history
9827_386	Restorationism (Christian primitivism) List of articles about Mormonism
9827_387	References
9827_388	Sources
9827_389	Leonard J. Arrington and Davis Bitton, The Mormon Experience: A History of the Latter-day Saints,
9827_390	(Urbana: 1992).
9827_391	Richard P. Howard, The Church Through the Years, (Herald House: 1992).
9827_392	Dallin H. Oaks and Marvin S. Hill, "Carthage Conspiracy", (University of Illinois Press).
9827_393	Further reading
9827_394	Matthew Bowman, The Mormon People: The Making of an American Faith Random House (January 24, 2012)
9827_395	Historiography
9827_396	Turner, John G. "More than a Curiosity: Mormonism and Contemporary Scholarship," Journal of
9827_397	Religion (April 2014) 94#2 pp 229–241. DOI: 10.1086/674956
9827_398	External links Official website of the LDS Church History Department
9827_399	Open Access Journal of Mormon History
9828_0	The J-2 is a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch
9828_1	vehicles. Built in the U.S. by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH2) and
9828_2	liquid oxygen (LOX) propellants, with each engine producing of thrust in vacuum. The engine's
9828_3	preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won
9828_4	approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February
9828_5	1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's
9828_6	performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type
9828_7	J-2T, which were cancelled after the conclusion of the Apollo program.
9828_8	The engine produced a specific impulse (Isp) of in a vacuum (or at sea level) and had a mass of
9828_9	approximately . Five J-2 engines were used on the Saturn V's S-II second stage, and one J-2 was
9828_10	used on the S-IVB upper stage used on both the Saturn IB and Saturn V. Proposals also existed to
9828_11	use various numbers of J-2 engines in the upper stages of an even larger rocket, the planned Nova.
9828_12	The J-2 was America's largest production LH2-fuelled rocket engine before the RS-25. A modernized
9828_13	version of the engine, the J-2X, was considered for use on the Earth Departure Stage of NASA's
9828_14	Space Shuttle replacement, the Space Launch System.
9828_15	Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be
9828_16	restarted once after shutdown when flown on the Saturn V S-IVB third stage. The first burn, lasting
9828_17	about two minutes, placed the Apollo spacecraft into a low Earth parking orbit. After the crew
9828_18	verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar
9828_19	injection, a 6.5 minute burn which accelerated the vehicle to a course for the Moon.
9828_20	Components
9828_21	Thrust chamber and gimbal system
9828_22	The J-2's thrust chamber assembly served as a mount for all engine components, and was composed of
9828_23	the thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark
9828_24	igniter.
9828_25	The thrust chamber was constructed of thick stainless steel tubes, stacked longitudinally and
9828_26	furnace-brazed to form a single unit. The chamber was bell-shaped with a 27.5:1 expansion area
9828_27	ratio for efficient operation at altitude, and was regeneratively cooled by the fuel. Fuel entered
9828_28	from a manifold, located midway between the thrust chamber throat and the exit, at a pressure of
9828_29	more than . In cooling the chamber, the fuel made a one-half pass downward through 180 tubes and
9828_30	was returned in a full pass up to the thrust chamber injector through 360 tubes. Once propellants
9828_31	passed through the injector, they were ignited by the augmented spark igniter and burned to impart
9828_32	a high velocity to the expelled combustion gases to produce thrust.
9828_33	The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed
9828_34	them in a manner that produced the most efficient combustion. 614 hollow oxidizer posts were
9828_35	machined to form an integral part of the injector, with fuel nozzles (each swaged to the face of
9828_36	the injector) threaded through and installed over the oxidizer posts in concentric rings. The
9828_37	injector face was porous, being formed from layers of stainless steel wire mesh, and was welded at
9828_38	its periphery to the injector body. The injector received LOX through the dome manifold and
9828_39	injected it through the oxidizer posts into the combustion area of the thrust chamber, while fuel
9828_40	was received from the upper fuel manifold in the thrust chamber and injected through the fuel
9828_41	orifices which were concentric with the oxidizer orifices. The propellants were injected uniformly
9828_42	to ensure satisfactory combustion. The injector and oxidizer dome assembly was located at the top
9828_43	of the thrust chamber. The dome provided a manifold for the distribution of the LOX to the injector
9828_44	and served as a mount for the gimbal bearing and the augmented spark igniter.
9828_45	The augmented spark igniter (ASI) was mounted to the injector face and provided the flame to ignite
9828_46	the propellants in the combustion chamber. When engine start was initiated, the spark exciters
9828_47	energized two spark plugs mounted in the side of the combustion chamber. Simultaneously, the
9828_48	control system started the initial flow of oxidizer and fuel to the spark igniter. As the oxidizer
9828_49	and fuel entered the combustion chamber of the ASI, they mixed and were ignited, with proper
9828_50	ignition being monitored by an ignition monitor mounted in the ASI. The ASI operated continuously
9828_51	during entire engine firing, was uncooled, and was capable of multiple reignitions under all
9828_52	environmental conditions.
9828_53	Thrust was transmitted through the gimbal (mounted to the injector and oxidizer dome assembly and
9828_54	the vehicle's thrust structure), which consisted of a compact, highly loaded () universal joint
9828_55	consisting of a spherical, socket-type bearing. This was covered with a Teflon/fiberglass coating
9828_56	that provided a dry, low-friction bearing surface. The gimbal included a lateral adjustment device
9828_57	for aligning the combustion chamber with the vehicle, so that, in addition to transmitting the
9828_58	thrust from the injector assembly to the vehicle thrust structure, the gimbal also provided a pivot
9828_59	bearing for deflection of the thrust vector, thus providing flight attitude control of the vehicle.
9828_60	Propellant Feed System
9828_61	The propellant feed system consists of separate fuel and oxidizer turbopumps (the bearings of which
9828_62	were lubricated by the fluid being pumped because the extremely low operating temperature of the
9828_63	engine precluded use of lubricants or other fluids), several valves (including the main fuel valve,
9828_64	main oxidizer valve, propellant utilization valve and fuel and oxidizer bleed valves), fuel and
9828_65	oxidizer flowmeters, and interconnecting lines.
9828_66	Fuel turbopump
9828_67	The fuel turbopump, mounted on the thrust chamber, was a turbine-driven, axial flow pumping unit
9828_68	consisting of an inducer, a seven-stage rotor, and a stator assembly. It was a high-speed pump
9828_69	operating at 27,000 rpm, and was designed to increase hydrogen pressure from (absolute) through
9828_70	high-pressure ducting at a flowrate which develops . Power for operating the turbopump was provided
9828_71	by a high-speed, two-stage turbine. Hot gas from the gas generator was routed to the turbine inlet
9828_72	manifold which distributed the gas to the inlet nozzles where it was expanded and directed at a
9828_73	high velocity into the first stage turbine wheel. After passing through the first stage turbine

9827_374

to give everything they had to this cause. However, when they failed to live the Law of

9827_375

Consecration, which was a promise to voluntarily give all their property to the community for equal

9827_376

distribution, Joseph Smith received a revelation that God would build Zion later, when the people

9827_377

were ready. Both the Community of Christ and the LDS Church have come to understand Zion as

9827_378

primarily a community way of life that is centered around Jesus Christ, but many people still

9827_379

believe that, when they are ready, they will be able to build the literal city of Zion as well.

9827_380

Further reading

9827_394

Matthew Bowman, The Mormon People: The Making of an American Faith Random House (January 24, 2012)

9827_395

Historiography

9827_396

Turner, John G. "More than a Curiosity: Mormonism and Contemporary Scholarship," Journal of

9827_397

Religion (April 2014) 94#2 pp 229–241. DOI: 10.1086/674956

9827_398

External links Official website of the LDS Church History Department

9827_399

Open Access Journal of Mormon History

9828_0

The J-2 is a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch

9828_1

vehicles. Built in the U.S. by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH2) and

9828_2

liquid oxygen (LOX) propellants, with each engine producing of thrust in vacuum. The engine's

9828_3

preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won

9828_4

approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February

9828_5

1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's

9828_6

performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type

9828_7

J-2T, which were cancelled after the conclusion of the Apollo program.

9828_8

The engine produced a specific impulse (Isp) of in a vacuum (or at sea level) and had a mass of

9828_9

approximately . Five J-2 engines were used on the Saturn V's S-II second stage, and one J-2 was

9828_10

used on the S-IVB upper stage used on both the Saturn IB and Saturn V. Proposals also existed to

9828_11

use various numbers of J-2 engines in the upper stages of an even larger rocket, the planned Nova.

9828_12

The J-2 was America's largest production LH2-fuelled rocket engine before the RS-25. A modernized

9828_13

version of the engine, the J-2X, was considered for use on the Earth Departure Stage of NASA's

9828_14

Space Shuttle replacement, the Space Launch System.

9828_15

Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be

9828_16

restarted once after shutdown when flown on the Saturn V S-IVB third stage. The first burn, lasting

9828_17

about two minutes, placed the Apollo spacecraft into a low Earth parking orbit. After the crew

9828_18

verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar

9828_19

injection, a 6.5 minute burn which accelerated the vehicle to a course for the Moon.

9828_20

Components

9828_21

Thrust chamber and gimbal system

9828_22

The J-2's thrust chamber assembly served as a mount for all engine components, and was composed of

9828_23

the thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark

9828_24

igniter.

9828_25

The thrust chamber was constructed of thick stainless steel tubes, stacked longitudinally and

9828_26

furnace-brazed to form a single unit. The chamber was bell-shaped with a 27.5:1 expansion area

9828_27

ratio for efficient operation at altitude, and was regeneratively cooled by the fuel. Fuel entered

9828_28

from a manifold, located midway between the thrust chamber throat and the exit, at a pressure of

9828_29

more than . In cooling the chamber, the fuel made a one-half pass downward through 180 tubes and

9828_30

was returned in a full pass up to the thrust chamber injector through 360 tubes. Once propellants

9828_31

passed through the injector, they were ignited by the augmented spark igniter and burned to impart

9828_32

a high velocity to the expelled combustion gases to produce thrust.

9828_33

The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed

9828_34

them in a manner that produced the most efficient combustion. 614 hollow oxidizer posts were

9828_35

machined to form an integral part of the injector, with fuel nozzles (each swaged to the face of

9828_36

the injector) threaded through and installed over the oxidizer posts in concentric rings. The

9828_37

injector face was porous, being formed from layers of stainless steel wire mesh, and was welded at

9828_38

its periphery to the injector body. The injector received LOX through the dome manifold and

9828_39

injected it through the oxidizer posts into the combustion area of the thrust chamber, while fuel

9828_40

was received from the upper fuel manifold in the thrust chamber and injected through the fuel

9828_41

orifices which were concentric with the oxidizer orifices. The propellants were injected uniformly

9828_42

to ensure satisfactory combustion. The injector and oxidizer dome assembly was located at the top

9828_43

of the thrust chamber. The dome provided a manifold for the distribution of the LOX to the injector

9828_44

and served as a mount for the gimbal bearing and the augmented spark igniter.

9828_45

The augmented spark igniter (ASI) was mounted to the injector face and provided the flame to ignite

9828_46

the propellants in the combustion chamber. When engine start was initiated, the spark exciters

9828_47

energized two spark plugs mounted in the side of the combustion chamber. Simultaneously, the

9828_48

control system started the initial flow of oxidizer and fuel to the spark igniter. As the oxidizer

9828_49

and fuel entered the combustion chamber of the ASI, they mixed and were ignited, with proper

9828_50

ignition being monitored by an ignition monitor mounted in the ASI. The ASI operated continuously

9828_51

during entire engine firing, was uncooled, and was capable of multiple reignitions under all

9828_52

environmental conditions.

9828_53

Thrust was transmitted through the gimbal (mounted to the injector and oxidizer dome assembly and

9828_54

the vehicle's thrust structure), which consisted of a compact, highly loaded () universal joint

9828_55

consisting of a spherical, socket-type bearing. This was covered with a Teflon/fiberglass coating

9828_56

that provided a dry, low-friction bearing surface. The gimbal included a lateral adjustment device

9828_57

for aligning the combustion chamber with the vehicle, so that, in addition to transmitting the

9828_58

thrust from the injector assembly to the vehicle thrust structure, the gimbal also provided a pivot

9828_59

bearing for deflection of the thrust vector, thus providing flight attitude control of the vehicle.

9828_60

Propellant Feed System

9828_61

The propellant feed system consists of separate fuel and oxidizer turbopumps (the bearings of which

9828_62

were lubricated by the fluid being pumped because the extremely low operating temperature of the

9828_63

engine precluded use of lubricants or other fluids), several valves (including the main fuel valve,

9828_64

main oxidizer valve, propellant utilization valve and fuel and oxidizer bleed valves), fuel and

9828_65

oxidizer flowmeters, and interconnecting lines.

9828_66

Fuel turbopump

9828_67

The fuel turbopump, mounted on the thrust chamber, was a turbine-driven, axial flow pumping unit

9828_68

consisting of an inducer, a seven-stage rotor, and a stator assembly. It was a high-speed pump

9828_69

operating at 27,000 rpm, and was designed to increase hydrogen pressure from (absolute) through

9828_70

high-pressure ducting at a flowrate which develops . Power for operating the turbopump was provided

9828_71

by a high-speed, two-stage turbine. Hot gas from the gas generator was routed to the turbine inlet

9828_72

manifold which distributed the gas to the inlet nozzles where it was expanded and directed at a

9828_73

high velocity into the first stage turbine wheel. After passing through the first stage turbine