Apollo 10

The Command Module Computer (CMC) is displaying its determination of their orbit in terms of its apogee and perigee. In metric, this is 190.0 by 187.2 kilometres and is regarded as essentially a circular orbit.

At the point of engine cut-off, the computer was displaying their inertial speed, their rate of vertical ascent and their altitude. They are travelling at 7,792 metres per second, ignoring the rotation of Earth, they were descending at around 30 centimetres per second (i.e. essentially flying parallel to the ground) and their altitude was 190.0 kilometres which means that they were inserted into orbit at their apogee.

The primary evaporator in the Environmental Control System began operation soon after lift-off but dried out after only a few minutes. The secondary cooling system was activated and functioned nominally. The primary evaporator was deactivated and was not reserviced with water until just prior to Lunar Orbit Insertion. It dried out again during the second lunar orbit. Just prior to entry, the evaporator was serviced again. During entry, it functioned normally.

This evaporator had dried out once during altitude chamber tests at KSC, and the cause was not determined. During later tests, the evaporator functioned satisfactorily.

After the mission, the spacecraft wiring and control circuits were checked. Continuity and resistance measurements were normal. Further tests of the system duplicated the inflight condition and revealed that the water control circuit operated intermittently. When a microswitch in this circuit opens, the water section of the environmental temperature control unit is activated and begins to supply water to the evaporator on demand.

A check of the switch assembly revealed that the actuator moved as little as 0.0008 inch beyond the point at which the switch should have opened. With changes in environment, the actuator travel was at times not sufficient to open the switch. Actuator rigging procedures were modified to assure proper overtravel on all subsequent missions.

Up to now, communications had been handled through a ground station near the launch site. With Apollo 10 heading towards the horizon for that site, communication moves to USNS Vanguard, a converted ship.

The Earth parking orbit around does not allow continuous communication with Houston, Mission Control. The orbit is low, when the spacecraft does come within range of a ground station, it is only above the ground stations horizon for a short period of time, even during passes that go directly overhead. For passes other than directly overhead, the period of contact is even less. Communications has been via the insertion tracking ship Vanguard. Following a one minute gap communications will be picked up by the tracking station on the Canaries.

When the propellant isolation valves in Command Module Reaction Control System B were opened about 10 hours prior to launch, the helium manifold pressure dropped from 44 to 37 psia. A pressure drop of this magnitude would be expected if the oxidizer burst disc was ruptured, allowing oxidizer to flow from the tank into the oxidizer manifold.

The isolation valve and the burst disc are redundant devices; therefore, a decision was made to proceed with the launch even though the disc was ruptured. The isolation valves were closed after orbital insertion. The engine valves were then opened by means of the reaction control heater circuits, and the oxidizer was vented from the manifold for 25 minutes. Afterward, the helium manifold pressure remained at 37 psia except for changes caused by thermal effects. When the isolation valves were opened just prior to system activation for entry, the helium manifold pressure dropped from 37 to 25 psia, confirming that the venting procedure had been effective and that the manifold was empty.

After the mission, the oxidizer and fuel burst discs were similar in physical appearance, indicating that the oxidizer burst disc had failed because of pressure.

Caution notes were added to the prelaunch checkout procedures for subsequent missions, in the places where the allowable limits on the burst disc (241 + or -16 psid in the flow direction and 10 psid in the reverse direction) could be exceeded. To allow early detection of any similar problem in the later missions, a leak check of the burst disc was added after Reaction Control System propellant servicing.

The space vehicle has now passed into range of the Tananarive tracking station on the island of Madagascar.

The VHF/AM transmitter-receiver is controlled by the VHF-AM controls on panel No.3 of the main display console (S43, 544, and S71). The Duplex/Off/Simplex switches activate the receivers and transmitters by applying 28-volt DC power. About 6 watts of power are required in these modes with the transmitter in standby and about 36 watts when keyed. Simplex A transmit and receive on 296.8 MHz for voice only.

Houston are trying to at least establish voice communications with the spacecraft through the Tananarive ground station.

Before the platform realignment could commence, the cover protecting the optics, which is located on the opposite side of the CM from the main hatch, was jettisoned to enable the spacecraft's optical instruments to view targets in deep space. The cover protects the external surfaces of the sextant and scanning telescope from the aerodynamic and thermal effects of the launch and ascent to Earth orbit.

Guidance and navigation are crucial to any journey, especially on the ballistic trajectory used to get to the Moon and back. The crew need the ability to point the spacecraft precisely, to well-known directions. The spacecraft carries a gyroscopically stabilised platform within the IMU (Inertial Measurement Unit) which remains fixed in attitude while the spacecraft moves about it, connected to it by three orthogonal gimbals. The only way to ensure that the platform is properly aligned is to compare its orientation to a fixed attitude reference - with the stars almost always used. Initially, the CSM (Command/Service Module)'s platform was aligned just before launch. Realignment, once in orbit, is achieved by using P52 [Program 52] on the CMC (Command Module Computer). With the spacecraft held in a steady attitude, the sextant is pointed at a specified star, marked on that star, then to a second star, where another mark is taken. The computer, knowing the attitude of the spacecraft (relative to its own idea of where the stars are), now has new values of where the stars are located. With these measurements, the amount of drift the platform has experienced can be calculated, and displayed as the amount of correction needed to move, or "torque" the gimbals to bring the platform back into accurate alignment. Known as "gyro torquing angles", they are displayed on the DSKY as "Noun 93". These angles are usually very small, and are expressed in thousandths of a degree.

The Command Module Pilot (CMP), John Young has completed these procedures.

Stafford is referring to the single eye patch sometimes used to aid the CMP when he is sighting through the CM optics

Stafford reference to "set that other one up in orb-rate" refers to the second FDAI (Flight director attitude indicator, aka "8-ball"). The FDAI can display the spacecraft's attitude in two different modes, Inertial or Orb-Rate. The FDAI displays the spacecraft's attitude relative to the celestial sphere (i.e. it normally shows the inertial attitude). For the FDAI to be operated in the Orb-Rate mode, a device know as ORDEAL (Orbital-Rate Drive Earth And Lunar) (panel 13) provides the correct drive signal to rotate the FDAI at a rate which also matches the orbital period. With the ORDEAL, the FDAI will display attitudes relative to the body's surface below (Earth or Moon).

Below is a description of the GDC taken from the Apollo Operations Handbook.

Gyro Display Coupler (GDC)

The purpose of the GDC is to provide a backup attitude reference system for accurately displaying the spacecraft position relative to a given set of reference axes. Spacecraft attitude errors can be displayed on an FDAI using the ASCP [Attitude Set Control Panel - In the lower left hand corner of Panel 1]-GDC difference. This error signal provides a means of aligning the attitude reference system to a fixed reference while monitoring the alignment process on the error needles; or it could be used in conjunction with manual manoeuvring of the spacecraft with the error needles representing fly-to commands.

The GDC can be configured for the following configurations:

• GDC Align - Provides a means of aligning the GDC to a given reference [i.e. the IMU]
• Euler - Computes total inertial attitude from body rate signal inputs.
• Non-Euler - Converts analogue body rate signals to digital body rate pulses.
• Entry (~.05 G) - Provides redundant outputs of attitude changes with respect to the roll stability axis.

GDC Configurations.

Panel switch positions necessary to obtain each particular GDC function are discussed below.

1. The GDC Align mode is used when aligning the GDC Euler angles (shafts) to the desired inertial reference selected by the ASCP thumbwheels (resolvers). This is done by interfacing the GDC resolvers with the ASCP resolvers (per axis) to generate error signals which are proportional to the sine of the difference between the resolver angles. When the GDC Align switch is pressed, these error signals are fed back to the GDC input to drive the GDC (ASCP resolver angular difference to zero. During the align operation all other inputs and functions for the GDC are inhibited. When the EMS ROLL [panel 1 bottom] switch is up and the GDC Align [panel 1 lower section] switch is pressed, the RSI pointer rotates (open loop) in response to yaw ASCP thumbwheel rotations.
2. In the Euler configuration, the GDC accepts pitch, yaw, and roll DC body rate signals from either gyro assembly and transforms them to Euler angles to be displayed on either FDAI ball. The GDC Euler angles also interface with the attitude set control panel (ASCP) to provide Euler angular errors, which are transformed to body angular errors for display on either FDAI attitude error indicators.
3. With the CMC ATT switch in the GDC position, pitch, yaw, and roll DC body rate signals from either gyro assembly are converted to digital body rate signals and sent to the G&N command module computer. Power is not only removed from both FDAI ball-drive circuits when this configuration is selected, but ASCP-generated errors are also removed.
4. In the entry mode (~.05 G), the GDC accepts yaw and roll DC rate signals from;
   1. Either gyro assembly, and computes roll attitude with respect to the stability axis to drive the RSI on the entry monitor system.
   2. Gyro assembly 1, and computes roll attitude with respect to the stability axis to drive either FDAI 1 or FDAI 2 in roll only.

Communications is now taken through the Honeysuckle Creek ground station in eastern Australia.