Threeaxis Hand Controller

FIGURE 4

, ia,c MCDONNELL

DAT. 26 November 1962 ke st. louis, missouri rao*

"tvistD-:__■feONI'IDMlSy «00«l Mercury Spacecraft

3-9 SPACECRAFT ENVIRONMENTAL CONTROL

3-9-1 ENVIRONMENTAL CONTROL SYSTEM - Control of suit and cabin environment shall be accomplished by means of an environmental control system (Drawing No. 1+5-83700). The environmental control system shall provide the following:

a. Environmental control, pressure 3uit (internal circuit)

b. Environmental control, cabin and equipment c. Cabin pressure relief d. Post-landing ventilation e. Cooling, prelaunch

3«9-1-1 DESCRIPTION - The environmental control system shall consist of a gaseous oxygen supply that shall furnish breathing, ventilation, and pressurization gas for the pressure suit and cabin. The environmental control system shall be designed to automatically control the environmental conditions within the pressure suit and cabin during all phases of the mission as described in Paragraph 1.1.1. Separate evaporative heat exchangers shall cool the suit circuit and cabin. Oxygen flowing from the suit circuit compressor shall pass through the carbon dioxide (CO2) and odor absorber, which shall be divided into individual sections that shall contain a supply of activated charcoal and lithium hydroxide (LiOH). The activated charcoal shall remove odor and the LiOH shall remove the CO2 from the gas flow. Filters shall be incorporated in the absorber to filter any charcoal or LiOH dust from the gas flow. An 02 partial pressure sensing system shall transmit a signal proportional to the amount of O2 partial pressure in the cabin to the indicator provided on the main instrument panel. A CO2 partial pressure sensing system shall transmit a signal proportional to the amount of CO2 partial pressure in the suit circuit to the indicator provided on the main instrument panel.

Moisture condensed in the pressure suit heat exchanger be absorbed and retained in a vinyl sponge. At timed intervals, the sponge shall be automatically compressed to force the condensate from the sponge to a condensate tank for storage. The sponge may also be compressed m»nin»ny by actuating a switch on the main instrument panel. The sponge ah«n be compressed by a piston actuated by oxygen pressure. Gas flow from the pressure suit passes through a solids trap which shall remove amy foreign matter, such as food particles, hair, nasal excretion, etc., from the suit circuit gas supply. The solids trap shall incorporate a relief feature to prevent the possibility of foreign matter blocking suit circuit flow.

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revised --report_6603-15A

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3«9-l-l DESCRIPTION - (Continued)

During re-entry, at approximately 17,000 feet, ambient air shall be directed into the cabin for cooling and ventilating. The equipment shall be as simple and passive in operation as practicable and shall provide the following:

a. Metabolic oxygen, pressurization and ventilation in the pressure suit and cabin.

b. Pressure suit ventilation for 12 hours of the post-landing phase.

c. A selectable cabin temperature between 50°P and 80°F during orbit.

d. Comfortable humidity-temperature level within the pressure suit during all phases of flight.

e. Cabron dioxide, moisture, odor, and solids removal.

f. Suit and cabin pressure regulation during all phases of flight.

g. A decompression feature for fire extinguishing.

h. Satisfactory operation in a weightless or high "g" environment.

i. A secondary oxygen supply. 3-9-1.2 OPERATIONAL SEQUENCE

3-9-1-2.1 PKKLAUNCH - During spacecraft prelaunch operation, the suit circuit (with facepiece dosed) and cabin be purged with oxygen from an external low pressure source. Freon 114 refrigerant shall be introduced into the pressure suit and cabin heat exchangers from an external source through the spacecraft umbilical connection to provide cooling for the suit and cabin. Cabin and suit temperatures shall be controlled by the ground crew by regulation of the Freon 114 flow to the heat exchangers. Prior to launch, the internal oxygen supply shall be activated by a ground crewman. Breathing gas from the Internal 02 supply shall be used by the astronaut during the count-down period. The Freon 114 refrigerant line shall be disconnected at spacecraft umbilical separation. A leakage check of the hatch seal shall be made by pressurizing the cabin to 19-7 psla via an external source through the cabin pressurization' fitting assembly.

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oat. 26 November 1?62 ,T> U0U|S> M(JMUW „„ _

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3-9-l'2.2 LAUNCH - During the launch operation, the cabin pressure relief valve will prevent the cabin-to-ambient differential pressure (aP) from exceeding 5.5 psig and shall maintain a cabin-to-ambient differential pressure (aP) of approximately 5.5 psig thereafter.

3.9.1.2.3 ORBITAL - The internal cabin temperatures during the orbital phase shall be dependent on the following:

a. Direct solar radiation absorbed at the outer surface of the vehicle.

b. Solar radiation reflected from the earth to the vehicle.

c. Direct radiation emitted from the earth to the vehicle.

d. Radiation emitted from the vehicle to the earth and space.

e. Internal heat generation from the astronaut and equipment.

f. Mass of the structure, insulation, equipment and furnishings.

Cabin temperature shall be regulated by adjustment of the cabin temperature valve. Cabin air shall be circulated by the cabin equipment fan which shall force the cabin gas through the equipment heat exchanger and around the electronic equipment. In the event of cabin pressure decay, repressurization shall be achieved by oxygen flow through the cabin pressure control valve. Pressure relief shall be afforded by the' cabin pressure relief and emergency decompression valve.

Oxygen shall be admitted from two primary oxygen bottles through pressure-reducing valves which shall drop the pressure from 75OO psig to 100 psig. The internal circuit pressure regulator shall supply the oxygen necessary to maintain approximately a 5 psla level during the orbital period. During ascent and descent, the suit pressure regulator shall also equalize suit internal and external pressure. A separate secondary bottle in parallel with the primary supply shall admit oxygen to the system through an oxygen pressure reducer which shall drop the pressure from 7500 psig to 80 psig. The cabin and suit circuits shall constitute redundant breathing and pressure sources, permitting the faceplece to be open or closed as desired by the astronaut.

The suit circuit compressor shall force gas through the pressure suit, solids trap, carbon dioxide and odor absorber, heat exchanger and water absorber. Pressure within the suit shall be maintained at 5 psia.

0ATE 26 November 1962

REVISED

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6603-15A

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In the event of failure of the main suit compressor, a back—up compressor shall be actuated automatically by the compressor differential pressure sensor. Should both compressors fail, the astronaut may breathe cabin atmosphere or utilize the emergency oxygen flow rate mode (see Paragraph 3'9'1'3'c)' In event of a meteoric collision causing depressurization of the cabin, the astronaut shall be able to continue by using the suit circuit for the full mission time at the normal oxygen usage rate or for one orbital cycle at the emergency flow rate of approximately 0.05 lb./min. In event of fire or buildup of toxic contaminants, the cabin may be decompressed.

3.9.1.2.4 RE-ENTRY - During re-entry, the environmental control system shall function as in the orbital sequence. Prior to re-entry initiation, cabin and cabin contents shall be cooled to as low a value as possible. Suit and cabin pressures shall remain at approximately 5 psia until an altitude of 25,000 feet is reached. At 17,000 feet, external air shall be automatically circulated through the suit circuit. In an emergency, a re-entry following a double failure of the recirculation system (with or without cabin depressurization) shall be accomplished using the emergency oxygen rate to provide breathing, ventilation, and pressurization of the suit. A reflective coating on the outer surface of the pressure suit will reduce radiant heat input.

3.9.1.2.5 POST-LANDING - Operational provisions shall be incorporated in the suit circuit for a 12-hour post-orbital period.. Ambient air shall be drawn into the suit circuit through a snorkel fitting, circulated, and exhausted overboard through a snorkel outlet.

' 3.9.1-3 OPERATIONAL MODES - The environmental control system shall operate automatically or manually in the following modes:

a. Cabin Mode - In this mode of operation the astronaut may have his suit faceplate open to the cabin environment. The cabin temperature shall be selected by the astronaut, by actuation of the knob located on the right-hand console (see Paragraph 3.8.9.6).

b. Suit Mode - In this mode the astronaut will have his suit faceplate closed and the cabin atmosphere will be excluded. The C0p content of the suit gas supply shall be maintained below 8 mm Hg. Comfortable combinations of temperature and humidity shall be selectable. Dual compressors shall be provided in the suit circuit, and the standby compressor shall be automatically switched on if the primary compressor fails. The astronaut also shall be able to switch in the standby compressor (see Paragraph 3.8.9.6). If the suit circuit falls, the emergency mode can be used. If the cabin system is operating normally when the suit circuit fails, the astronaut may open the faceplate instead of actuating the emergency mode.

26 November 1062

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3-9-1.3 OPERATIONAL MODES - (Continued)

c. Emergency Mode - In this mode of operation em automatic and/or manual emergency oxygen rate capability shall be provided. The emergency oxygen rate may be used during loss of cabin pressurization or during failure of the closed environmental control system. This oxygen shall be available for use in the suit mode (b) described above. This system shall be used through the suit by a direct open oxygen system in which expired oxygen is discharged by being dumped into the cabin and then overboard . Provision shall be made to pexmit the use of remaining primary oxygen supply for this mode; however, special provision shall be made to prevent loss of oxygen to the cabin system if the cabin system fails.

3-9-1.1+ ENVIRONMENTAL CONTROL SYSTEM WARNING INDICATION - Amber warn ing lights with accompanying audio tones shall be provided on the warning light portion of the main instrument panel (see Paragraph 3.8.9.4.1) for indication of the following:

O2 EMERGENCY

O2 QUAN

EXCESS H2O

C02 PRESS

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3.10 STABILIZATION AND CONTROL SUBSYSTEM - The stabilization and control subsystem shall consist of the automatic stabilization and control system, the horizon scanner system, and the reaction control system. The launch trajectory control and guidance shall be considered an integral part of the launch vehicle system and shall not be the responsibility of the spacecraft contractor.

3.10.1 AUTOMATIC STABILIZATION AND CONTROL SYSTEM - The automatic stabilization and control system (ASCS) as defined in Drawing No. 45-87700 (see Appendix I-C, Item 5) shall provide automatic stabilization and orientation of the spacecraft from time of separation from the booster-adapter until landing parachute deployment in accordance with the various phases of the mission. The ASCS shall supply output signals for display, recording and telemetering of three-axis attitude information, a discrete signal at 0.05g longitudinal acceleration during re-entry, and attitude signal sectors for use in the spacecraft retrograde firing interlock circuit. Associated equipment consisting of the horizon scanners, reaction controls, communications system telemetry, devices for display of the spacecraft attitude, and devices for generating spacecraft signals for discrete mission events, shall be utilized by the ASCS. The expenditure of propellant shall be minimized by the design of the control system.

3.10.1.1 MODES OF OPERATION - The ASCS shall have four modes of automatic operation: damping, orientation, attitude-hold, and re-entry. The ASCS shall also contain switching to allow alternate manual fly-by-wire and auxiliary damping modes. In the fly-by-wire mode, the automatic reaction control nozzles shall be controllable by the astronaut through limit switches actuated by stick controller motion. The auxiliary damping mode shall provide rate damping only and shall disengage the automatic and fly-by-wire functions. Other switching is available for manual selection of an orbit pitch attitude of either the preset retrograde attitude of 34° or the preset post-retrograde attitude of 0°, and for orbital rate pitch precession at astronaut discretion. Other than override controls, no provisions shall be made for manual stepping of the ASCS automatic sequencing. Provisions shall be made for an audio tone and light to be activated in the event of loss of attitude-hold mode (i.e., high thrust firing command) during orbit.

3.10.1.2 SEQUENCE OF OPERATION - The following general sequence of operation compatible with spacecraft sequence shall be provided by the ASCS.

a. Rate damping in early abort cases.

b. Rate damping, turn-around, and spacecraft orientation to the orbit attitude of 34° (blunt end up) In later aborts or after spacecraft-adapter separation in normal orbital missions.

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revised st. louis, missouri report raoe

revised st. louis, missouri mooel Mercury Spacecraft

3.10.1.2 SEQUENCE OF OPERATION - (Continued)

c. Orientation during orbital flight through retrograde rocket firing as follows:

1. Orientation with respect to the local earth vertical (such that the astronaut's head would be up).

2. Provide required spacecraft orientation prior to retrograde rocket firing.

3. Hold prescribed 34° (+5°) retrograde pitch attitude during retrograde rocket firing.

d. Switching to re-entry mode at retrograde assembly Jettison providing spacecraft orientation to a prescribed re-entry pitch attitude of 0° (+5°) following retrograde rocket firing. ~

e. Hold 0° (+5°) re-entry pitch attitude until 0.05g acceleration is sensed.

f. Switching to rate damper mode at longitudinal acceleration (from drag buildup) of 0.05g and providing a steady roll rate of approximately 10 to 12 degrees per second thereafter until disengagement.

g. Disengagement when landing chute deploys.

' The ASCS shall Include, in addition to amplifier-calibrator (computing) equipment, pitch, roll and yav rate gyros, vertical and directional attitude gyros, and a longitudinal accelerameter. The rate gyros shall sense spacecraft rotational rates, and the longitudinal accelerameter shall sense 0.05g longitudinal acceleration Initiation of the re-entry mode. The attitude gyros, with signal Inputs from the horizon scanners and slaving computation performed in the amp-cal, shall sense pitch, roll and yaw attitudes for an attitude reference system. The pitch and roll outputs of the horizon scanners shall be utilized to process the gyros such that their spin axes shall be maintained In the properly erected position relative to the moving local vertical axis. Prior to launch, both the vertical and directional gyros shall be torqued so as to erect their spin axes to any desired orientation relative to the launch trajectory. During the climb phase of the mission after tower separation, the vertical gyro spin axis shall be erected to the horizon scanners. After tover separation the vertical and directional gyros and horizon scanners shall function as shown In Figure 6, Page 66. At 0.05g, the horizon scanners shall be de-energized and the spacecraft pitch and yaw angular rates shall be maintained at a value which will Impose tolerable acceleration levels on the

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3-10.1.2 SEQUENCE OF OPERATION - (Continued)

astronaut and equipment. At the same time, a steady-state roll rate of approximately 10 to 12 degrees per second shall be established and maintained until disengagement of the ASCS at main landing parachute deployment.

3-10.2 HORIZON SCANNER SYSTEM - A horizon scanner system (Drawing No.

45-57702) shall be provided for sensing roll and pitch attitude reference for the ASCS. The horizon scanner system shall consist of two scanner units, one unit aligned to the spacecraft pitch axis and one unit aligned to the spacecraft roll axis. The scanner assemblies shall be mounted on the structure within the antenna assembly, and shall provide a llfl-degree conical scan of the horizon through a rotating prism located ahead of the scanner lens. The prism shall rotate at a speed of approximately 30 revolutions per second. Each scanner unit shall receive A.C. power inputs through the spacecraft A.C. power system and shall supply D.C. output signals of the required polarity to provide roll and/or pitch signals up to a maximum of 35 degrees for torquing the attitude gyros in the ASCS. Yaw sensing shall be achieved through torquing of vertical and directional gyros of the ASCS by horizon scanner roll signal inputs. Pitch and roll sensing shall possess sufficient accuracy to enable the astronaut to orient the spacecraft within +5° of the orbital attitude. The scanners shall run only when the gyro control switch on the main instrument panel is the the GYRO SLAVE position. In a normal mission the scanners shall be energized at time zero and shall function as illustrated in the Functional Profile, Figure 6 Page 66.

Individual thermostatically controlled blanket heaters (Drawing No. U5-78071) shall be provided for each horizon scanner to maintain the required scanner operating temperature. These heaters shall be energized when power is applied to the 24-volt D.C. bus and shall function continuously through antenna assembly jettison.

A protective fiberglas cover assembly (Drawing No. 45-31064) shall be provided for the roll scanner. The cover shall protect the scanner glass from blast erosion from the escape rockets, and shall be automatically ejected to permit horizon scanning at tower Jettison plus four seconds (see Paragraph 3*5*6).

3.10.3 REACTION CONTROL SYSTEM - The reaction control system

(Drawing No. 45-61700) (see Appendix I-C, Item 6) shall consist of an automatic control subsystem and a manual control subsystem, as depicted in Figure 5* Page 55* The reaction control system shall provide control of the spacecraft in the roll, pitch and yaw axes. This system shall be a pressure-fed, monopropellant/catalyst bed design, incorporating right angle firing exhaust nozzles which shall produce thrust through decomposition of hydrogen peroxide (H2O2). Minima], translation motions may result upon application of reaction control thrust.

date 26 November 1962

revised revised '

RASE 53

REPORT_66OT-1SA

mooel Mercury Spacecraft

3.10.3.1 AUTOMATIC CONTROL SUBSYSTEM - The automatic control subsystem shall basically consist of a pressurization system, fuel distribution system, and 12 thrust chamber assemblies. Each thrust chamber assembly shall consist of a solenoid valve, heat barrier, and thrust chamber. The fuel supply shall be unstabilized HgOg contained Inside a flexible bladder which, in turn, shall be contained in .a half-toroidal tank. This system shall function automatically in conjunction with the automatic stabilization and control system. A pressure transducer In the pressurization system shall provide a means of monitoring (by the pressure-versus-volume method) the percentage of fuel present in the bladder. Sufficient fuel and pressurization gas shall be provided to maintain damper operation until main parachute deployment, at which time the fuel shall be Jettisoned.

3.10.3.2 MANUAL CONTROL SUBSYSTEM - The manual control subsystem shall consist of a pressurization system, a fuel distribution system and six thrust chamber assemblies. The yaw, pitch, and roll control each consist of a pair of thrust chamber assemblies with a single proportional control propellant valve. The pressurization portion of the manual control subsystem shall be identical to the corresponding portion of the automatic subsystem. The manual subsystem shall have a smaller capacity for fuel than the automatic subsystem. The manual subsystem shall be controlled by the astronaut by means of three-axis hand controller (see Paragraph 3.8.8.2.1), and shall be capable of overcoming the disturbance torque resulting from' firing the retrograde rockets.

3.10.3.3 OPERATION - High-pressure nitrogen shall be utilized to pressurize the fuel tanks. The high-pressure (3000 pslg)

gas shall pass through a filter and manual shutoff valve to a pressure regulator (which shall reduce the pressure to 480 pslg), through a check ' valve, and finally surround and compress the flexible bladder of the torus tank. The pressure shall force fuel out of the bladder through the perforated tube downstream into the lines and valves. The manual push-pull shutoff valves, which allow fuel to be available at the solenoid valves, shall provide a means of individual system isolation and shutoff. The shutoff valves shall be vented overboard through a line system (Drawing No. 45-62075) to reduce the possibility of fire due to fuel leakage. Upon receiving a 24-volt D.C. signal from the ASCS, the appropriate solenoid valve shall open. The fuel shall then pass into the corresponding thrust chamber where it shall be decomposed, providing the following thrust levels for operation with the ASCS:

a. High thrust level of 24 pounds for pitch and yaw axes and six pounds for the roll axis.

b. Low thrust level of one pound for all three axes.

These thrust levels shall be available in discrete, short-time periods as controlled by the ASCS.

0ATK 26 November 1962

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6603-15A

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3.10.3.3 OPERATION - (Continued)

A pressure transducer in the pressurization system shall provide a means of monitoring the percentage of fuel present in the bladder. The Internal transfer tube shall guarantee uninterrupted and total fuel flow. The external transfer tube shall assure that no nitrogen shall be trapped during propellant filling.

The manual control subsystem shall provide proportional control and thrust levels between four and 2k pounds for pitch and yaw axes and between one and six pounds for roll axis. These thrust outputs shall be controlled from the hand controller through direct stick control.

3.10.3.4 TANKS - The pressurant tanks for each system shall be located in the cabin and shall be of spherical fiberglas construction.

These tanks shall store pressurization gas at 3000 psig. The fuel tanks shall be a half-toroidal configuration contoured to mount on the aft pressure bulkhead between the bulkhead and the heat shield. The fuel tanks shall be constructed of aluminum, insulated to provide temperature control, and incorporate a flexible plastic bladder to provide pressure for positive expulsion of the fuel. Provisions for in-flight Jettisoning of fuel at main parachute deployment shall be Incorporated. The manual and automatic fuel tanks shall be interconnected by a line system which permits use of fuel remaining in either tank in the system normally supplied by the other tank. A manual shutoff valve shall be incorporated in the interconnecting line for optimum utilization of available fuel at the discretion of the astronaut. Check valves shall be Installed in the line to prevent fuel interflow between tanks. The interconnect system shall also provide the method by which the fuel remaining in the manual tank at main parachute deployment shall be Jettisoned.

Provisions for an auxiliary fuel tank of 15 pounds capacity shall be incorporated on the spacecraft. This tank shall be of the same configuration as the automatic and manual fuel tanks except for size, and shall be in parallel with the automatic control system. Pressurization shall be provided by either the manual or automatic control system pressurant as selected by the astronaut. Pressurization shall be effected through squib valves, which shall be Initiated by a switch on the left-hand console. The squib valves shall remain de-energlzed until retrograde to insure availability of auxiliary fuel for retrograde and re-entry stabilization.

nevisco nevisco

26 NOVEMBER 1962

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