PART III Midcourse Guidance–Jack Fisher

The analysis of midcourse maneuver requirements had been previously well defined by JPL for their Ranger and Mariner missions and was primarily dedicated to correcting injection errors resulting from launch vehicle performance dispersions. The Surveyor mission introduced another level of complexity to the problem of midcourse guidance with the requirement for a lunar soft landing at a pre-selected location. For the selected lunar trajectories the spacecraft will approach the moon at a velocity as great as 2700 meters/second. To achieve a soft landing this velocity would be reduced to zero in two stages. The spacecraft design utilized a solid propellant rocket motor to remove the bulk of this speed before the start of the second stage terminal descent utilizing the liquid propellant vernier engines. The start of the terminal descent is signaled by an Altitude Marking Radar signal at a slant range of 60 miles above the lunar surface at which point the propulsion system operation is enabled after a ground commanded time delay. The remainder of the terminal descent sequence is fully automatic with no means to alter by ground command. Thus the only means to control terminal descent parameters after the midcourse maneuver is the ground commandable time delay that can range only from 0 to 20 seconds. This emphasizes the importance of the midcourse maneuver in ensuring a successful terminal descent.

Unbraked lunar approach speeds can vary by as much as 80 meters per second over a given launch period. Since the solid propellant loading cannot be modified during a launch period the midcourse maneuver can be utilized to make adjustments in the main retro burnout speed so that a fixed solid propellant loading would prove to be appropriate.

JPL’s Surveyor specification required the capability to perform midcourse maneuvers and a controlled soft landing on the lunar surface at a designated location. Hughes had the tasks of performing the analysis to determine the maneuver requirements, developing the software to be used during real-time mission operations, and providing the personnel to conduct the operations in JPL’s SFOF.

Requirements for the propulsion subsystem include attitude stabilization during the main retro firing and terminal descent and led to a system design with three throttleable vernier engines that would also be used for the midcourse maneuver. Thus the requirements for the midcourse maneuver must take into account the predicted vernier propellant requirements for the remainder of the mission.

The first step in planning a midcourse maneuver is to select an execution time. Surveyor specifications state the maneuver could be performed as early as 4 hours and as late as 40 hours after injection. The nominal correction time is 15 hours after injection for direct ascent and 20 hours for parking orbit missions. The nominal times are based upon the desire for Goldstone DSIF visibility for the maneuver and most desirable would be in the middle of the view period so as to allow immediate assessment of the executed maneuver. Two other variables are the orbit determination accuracy on which the maneuver will be based and the maneuver ∆V magnitude. The first argues for allowing sufficient time to for accurate orbit determination and the second for accomplishing the maneuver earlier as the required ∆V for a given correction increases with decreasing distance to the moon.

With the selection of maneuver time the position of the spacecraft is fixed and there remains the capability of adjusting the three velocity components. Two velocity components can be used to adjust the landing location and these define what is called the critical plane. The other velocity component, normal to the critical plane, can adjust the time of flight or other related parameters such as the unbraked lunar impact velocity. These two corrections are almost independent and to the first order can be determined separately.

The first step in the midcourse analysis is to determine the maneuver required in the critical plane to place the landing site at the desired location. This is accomplished by determining the sensitivity of the landing site location to velocity increments in the critical plane and using this data to formulate a covariance matrix can be used to determine the required maneuver. With the critical plane component known the velocity component normal to the plane can be determined by considering the applicable constraints. These include time of flight considering Goldstone visibility limits, the main retro burnout velocity constraints and the vernier propellant requirements for the remainder of the mission. Further the consumption of vernier propellant for the midcourse maneuver lowers the spacecraft weight and results in a greater main retro ∆V and a lesser main retro burnout velocity.

The software to determine midcourse maneuver requirements during mission operations was developed by Mal Meredith, John Ribarich and Len Davids and is described in the document, Midcourse and Terminal Guidance Operations Programs (Hughes SSD 4051R) published in April 1964. This program will determine a velocity increment at a specified time that will result in a landing at the desired landing site and ensure that:

1)   Main retro burnout velocity constrained by the successful operation of the radar altimeter and Doppler velocity sensor (RADVS) and the desired descent profile.

2)   The remaining fuel must be sufficient for the main retro and terminal descent operations including a reserve for possible dispersions.

3)   The unbraked arrival incidence angle must be less than a defined maximum value that will allow the altitude marking radar to operate satisfactorily.

4)   The arrival time must satisfy Goldstone DSIF visibility requirements.

The midcourse and terminal guidance program also performs a number of other calculations including required attitude maneuvers to implement the midcourse and terminal maneuvers, determine potential errors in the terminal conditions, vernier propellant requirements for the remainder of the mission, potential alternate landing sites, and investigate alternate mission scenarios including possible hard landings or flybys.

Midcourse Guidance–Flight Experience

For six of the seven Surveyor missions a midcourse maneuver was successfully accomplished. The maneuver for the Surveyor II mission resulted in a vernier engine failure that caused the spacecraft to tumble . However, the analysis for this planned maneuver is included in this discussion.

The performance of the Atlas-Centaur launch vehicle was outstanding. Not only were all Surveyors successfully launched, but the injection accuracy was remarkable. If only the injection errors were to be corrected the average midcourse velocity increment would have been 3.1 meters per second for the seven missions. For the last 3 missions the increment was only 1.1 meters per second.

Actual maneuver planning was more comprehensive taking into landing site accuracy and concern for time of flight, main retro burnout speed and the vernier propellant required for the terminal descent phase.

Landing accuracy upon lunar trajectory injection was expected to be a 99% circle 50 kilometers in radius. For later missions this was reduced to a 30-kilometer radius. As could be expected there would be quite a variation in lunar terrain in an area of this size some of which would not be suitable for a soft landing. It was therefore important in maneuver planning to attempt to reduce this uncertainty. Also as Lunar Orbiter photos became available they were used to select areas of suitable terrain. For example, with the Surveyor I mission the landing site was biased about one degree to the north to avoid several craters. The spacecraft actually landed about 15 kilometers from this aim point. The Surveyor I non-critical velocity component was selected to be 20 meters per second to reduce the main retro burnout velocity by about 100 feet per second. This maneuver component also provided close to the maximum vernier propellant margin and a time of flight that resulted in landing in the middle of the Goldstone view period.

For the seven missions the average non-critical plane maneuver was 9.2 meters/second and the critical plane component was 3.3 m/s, but this includes a maneuver of 11 m/s for Surveyor VII. For this mission the landing site was moved from the equatorial region to 41 degrees south latitude to allow scientific exploration of the lunar highlands region. Maneuver times for six of the seven missions were between 16 and 21 hours after launch in the first Goldstone pass. For Surveyor IV the maneuver was conducted during the second Goldstone pass at 38.5 hours in order to take advantage of the small maneuver requirement and improve landing accuracy.

The Surveyor V mission provided the most challenging midcourse guidance situation. The planned and executed maneuver at 17 hours after launch was 0.5 m/s in the critical plane and 14 m/s in the non-critical direction for a total maneuver of 14.01 m/s. It was determined after this maneuver that helium regulator valve was stuck open and the helium pressurant required for operation of the vernier propulsion system was leaking. In an attempt to reseat the valve three more maneuvers were conducted over the next three hours—none were successful in stopping the leak. A fifth maneuver was performed at 24 hours after launch to reduce main retro burnout velocity and maximize usable propellant in the terminal phase by increasing gas volume in the propellant tanks. A sixth maneuver of 5.3 m/s was performed at 39 hours after launch totadjust spacecraft weight and burnout velocity and further increase gas volume to optimal values, and correct a 267-km residual miss to move the impact point from an area of mountainous terrain to the desired landing site. The terminal descent and landing were successful.

Part II: Free Flight Trajectory Design–Jack Fisher

According to the terms of the Surveyor contract Hughes was responsible for Surveyor mission design and mission flight path operations. This was not JPL’s preference, but NASA headquarters dictated this due to JPL’s heavy involvement at the time with the Ranger and Mariner missions. Hughes had no experience with lunar trajectories prior to the start of the Surveyor contract. JPL’s RFP provided reference trajectory data for use in the design of the spacecraft so there was no need to become involved in this area during the proposal effort. The Hughes effort was rightly directly solely towards the spacecraft design. Mal Meredith had requested permission during the proposal to gain some experience in this area, but his request was denied. At the beginning of the Surveyor contract there was evidently no background or experience in this realm. When I first became involved with Surveyor in early 1962, about one year after the contract start, the engineers that I looked to for guidance were Mal Meredith and Paul Wong. At that time responsibility for Surveyor trajectory design resided outside the Surveyor project within Eli Botkin’s section in Ed Marriott’s Engineering Mechanics and Preliminary Design Department.

In May 1962 JPL requested that the Hughes Surveyor program office provide several Hughes engineers to take up residence at JPL to learn more about JPL mission design and operations for the Ranger lunar mission and subsequently apply this knowledge to Surveyor. The areas of concern were trajectory analysis, midcourse guidance and orbit determination. The engineers selected were myself, John Ribarich and Mike Horstein.

The purpose of my assignment at JPL was to design the lunar trajectories for Ranger 6 that was scheduled for launch in January 1963. The Ranger mission was intended to obtain pictures of the lunar surface prior to impact on the lunar surface. I spent about six months working at JPL primarily with Bill Kirhofer and Vic Clarke learning the ins and outs of lunar trajectories and mission operations. At that time Vic Clarke was conceded to be the leading expert for lunar and planetary trajectories in the country.

As a part of this assignment I took part in mission operations for Ranger 5 that was launched on October 18, 1962. A malfunction after 15 minutes of operation caused the irreversible transfer of power from the solar panels to the battery. After about 9 hours of operation the battery was depleted and operation of the spacecraft ceased. This failure resulted in the replacement of JPL’s Ranger project manager and a postponement of the Ranger 6 mission that was not launched until January 1964. Due to this delay I terminated my assignment at JPL in early 1963 and returned to Culver City.

Upon my return to Hughes I wrote a detailed description of JPL lunar trajectory operations. Shortly thereafter Mal Meredith convinced me to join Surveyor project team and assume responsibility for the design of Surveyor lunar trajectories under Bill Grayer in the Guidance and Trajectory Department in Jim Cloud’s Systems Engineering and Analysis Laboratory. My experience at JPL formed the basis for the Hughes Surveyor lunar trajectory efforts.

Hughes tasks included selecting the Surveyor lunar transfer trajectories, specifying trajectory parameters to General Dynamics for the Centaur guidance equations, developing the logic and software for midcourse guidance as well as terminal descent and guidance. The major thrust was to provide the necessary data for the overall design of the Surveyor spacecraft. Further, for mission operations to be conducted in JPL’s Space Flight Operations Facility (SFOF) Hughes would have the responsibility for staffing and the managing the Flight Path Analysis and Command (FPAC) group and the Spacecraft Analysis and Command (SPAC) group.

Within Bill Grayer’s Guidance and Trajectory Department I assumed responsibility for the free flight trajectory design, John Ribarich for midcourse guidance analysis, and Len Davids for terminal descent and guidance. Mal Meredith assumed the role of “systems engineer” to assure that our efforts were properly coordinated. Mal later was assigned the role of FPAC director and played a major role in planning and conducting Surveyor flight operations.

The successful design of the Surveyor lunar trajectories required the establishment of the relevant requirements and constraints. For Surveyor these were:

• The launch vehicle is the Atlas Centaur utilizing either the direct ascent or the parking orbit operational mode.

• For the parking orbit mode the minimum coast time is 116 seconds and the maximum coast time is 25 minutes.

• Launch will take place from Pad 36 of the Air Force Eastern Test Range (now referred to as the Kennedy Space Center).

• Launch azimuths from Pad 36 are restricted to the range from 80 to 115 degrees east of true north for direct ascent and 78 to 115 degrees for parking orbit.

• The Centaur payload capability is limited by the requirement for a 3-sigma propellant reserve.

• Arrival at the moon shall be with Goldstone visibility a minimum of 2 hours prior to and 3 hours following unbraked impact with a time of flight near 66 hours. Goldstone is the DSIF station located near Barstow, California.

• For the initial Surveyor missions the landing sites were selected to be within the Apollo landing zone of plus or minus 45 degrees lunar longitude and plus or minus 5 degrees lunar latitude. Note: the first six Surveyor missions with four successful landings satisfied Apollo requirements so Surveyor VII successfully landed in the lunar highlands at 40 degrees south lunar latitude to provide science data.

• For direct ascent trajectories lunar landing shall be between 20 hours before the morning terminator and 72 hours before the evening terminator.

• For parking orbit trajectories lunar landing shall be between 20 hours after the morning terminator and 150 hours before the evening terminator.

• The unbraked lunar impact speed shall be between 2615 and 2692 meters per second for direct ascent missions and 2600 and 2690 meters per second for parking orbit missions.

• The unbraked flight path angle at impact measured from the vertical was constrained to 25 degrees for the early Surveyor missions and 45 degrees for the later missions.

The establishment of lunar trajectories can proceed with several levels of precision. Many important characteristics can be determined with simple earth-centered conics including establishment of the dates on which it was possible to launch towards the moon and the times on those days during which it was possible to launch. Given that the moon orbits the earth with a period of 27.3 days launch opportunities should repeat on a near-monthly basis. In JPL terminology the times during a day during which a launch is possible are called the launch window while the days during a month on which a launch is possible are called the launch period. Once the launch opportunities are determined the trajectory characteristics can be established and made available for the detailed design of the spacecraft and the planning of mission operations.

Of special interest are the conditions at the moon for initiation of the terminal descent. To determine the lunar trajectory characteristics it is necessary to add a further degree of complexity to the analysis. The so-called patched conic approach requires transferring the trajectory from an earth-centered phase to a moon-centered phase at the point where the earth’s and moon’s gravitational attractions are equal. At this point an earth-centered ellipse becomes a moon-centered hyperbola and the lunar approach characteristics can be determined.

Most of the studies conducted were based upon the patched-conic approach. In 1965 Ron Gillett and I published our studies of Surveyor lunar trajectories. Direct ascent launch opportunities from January 1966 to December 1968 and parking orbit launch opportunities from January 1967 to January 1969 were presented in two separate reports.

More precise simulations of lunar trajectories were required to establish the accuracy of the simpler solutions and also to provide data to General Dynamics for the determination of the Centaur guidance equations and firing tables for each mission. JPL provided Hughes with a all-encompassing trajectory simulation that satisfied these accuracy requirements. However, the running time on the computers of that time was prohibitive. Fortunately, Paul Wong with Gladys’ Perkins programming assistance was able to create a simpler simulation of sufficient accuracy based upon Encke’s method. This simulation took into account several perturbing factors including the Earth’s oblateness and the Sun’s gravitational attraction. The accuracy of this program was established by comparing the results with those of JPL’s program. Initially there were serious discrepancies between the two programs, but these were eliminated with the use of double precision integration.

It was necessary to specify to General Dynamics the launch parameters that would result in a successful mission based upon the desired mission parameters at the moon.   These were the landing site location determined by lunar latitude and longitude and the arrival time and arrival speed. For launch on a given day with parking orbit ascent the four factors that can be varied to achieve the desired landing conditions are launch time, launch azimuth, parking orbit coast time and the injection energy defined by JPL in terms of C₃, a measure of the energy remaining after escape from the Earth’s gravitational field. For direct ascent trajectories the four launch parameters were launch time, launch azimuth, flight path angle at injection and injection energy, C₃. Trajectories were calculated varying the launch parameters to determine the impact on the lunar arrival conditions. A covariance matrix was created that related the launch conditions to the arrival conditions so that launch parameters resulting in the desired landing conditions could be determined. These trajectory parameters were then specified to JPL and GDC. A post-injection standard trajectories report was generated for each mission—these reports were written by Carl Winkelman for the first two Surveyors and Stan Dunn for the last five.

Fly Me to the Moon: The Surveyor Mission and Mission Design–Jack Fisher

The Surveyor mission as defined by NASA and JPL in 1960 was to soft land a package of instruments on the moon so that this payload would operate after landing. In January 1961 when Hughes was awarded the contract for Surveyor no spacecraft had safely landed on the moon. The first attempted lunar missions took place in 1958; four Pioneer missions by the United States attempted to orbit the moon and three Soviet Luna missions attempted to impact the moon—none were successful. The Soviet Luna 2 was the first spacecraft to impact the moon in September 1959 followed by Luna 3 in October that returned images of the far side of the moon. It was against this background that JPL and Hughes Aircraft set out to design a spacecraft that could achieve this goal.

This Surveyor mission description will be documented in four separately posted reports: 1. Surveyor Mission Description, 2. Free Flight Trajectory Design, 3. Midcourse Guidance and 4. Terminal Guidance and Descent.

1. Surveyor Mission Description

All seven Surveyor missions were successfully launched by an Atlas-Centaur utilizing Complex 36 of the Kennedy Space Center. Two different launch modes were used—direct ascent using a continuous Centaur burn and a parking orbit ascent during which the Centaur shuts down and then restarts after a coast period. When NASA’s Lewis Research Center (now Glenn) took over the Centaur development program it was decided to delay development of the parking orbit capability until the other Centaur issues had been resolved. The initial Centaur design wasn’t able to provide the required propellant settling during the coast period to allow an engine restart. For the historical record three launches used the direct ascent mode: Surveyor I (5/30/66), Surveyor II (9/20/66), and Surveyor IV (7/14/67); and four used the parking orbit mode: Surveyor III (4/17/67), Surveyor V (9/8/67), Surveyor VI (11/7/67) and Surveyor VII (1/7/68).

Note: The following edited mission summary was originally published in Hughes SSD 56028R Surveyor Trajectory Characteristics, December 1965.

The Atlas-Centaur boosts the spacecraft to into a lunar transfer trajectory that will reach the moon in approximately 66 hours. Following separation from the spacecraft the Centaur is turned 180 degrees and the remaining propellant is expended so that the vehicle will not impact the moon or interfere with Surveyor operations. The Surveyor after separation deploys the landing legs, omni antenna, and solar panel to its cruise position. Following these deployments the attitude jets stabilize any separation tip-off rates and a sun acquisition sequence is initiated.

Atlas Centaur Surveyor Parking Orbit Ascent Trajectory

Approximately 6 hours after launch, a command is given to acquire the star Canopus, by means of a roll maneuver, while maintaining sun lock. After Canopus acquisition is signaled by the spacecraft, a verification procedure is performed to ensure the sensor has locked on Canopus and not another star or planet. At this point Surveyor is fully attitude stabilized.

During the initial coast phase engineering telemetry is gathered to assess the health of the spacecraft and tracking data is utilized to accurately determine its orbit. Nominally at 15 to 20 hours after launch, with visibility from the Goldstone DSIF, a midcourse maneuver is performed using the three throttleable vernier engines. This maneuver corrects Atlas-Centaur injection errors, adjusts the location of the landing site if desired, and assures that the initial conditions for terminal descent provide high probability for soft landing. On command, the spacecraft performs a series of attitude changes that aligns the thrust axis in the required direction. The three vernier engines are then ignited and burn at a fixed acceleration level for a commanded time to provide the desired velocity increment.

Surveyor Lunar Transit Trajectory

Following the completion of the midcourse maneuver, the spacecraft is returned to the cruise attitude locked on the sun and Canopus. Engineering telemetry and tracking data are gathered by the DSIF and furnished to the SFOF for analysis of spacecraft status. Further orbit determination allows evaluation of the midcourse maneuver and computation of final terminal maneuver parameters.

As the spacecraft, in view of the Goldstone DSIF, approaches the moon 66 hours after launch, the thrust axis is aligned with the velocity vector. The high-gain antenna is pointed at the earth if approach television pictures of the lunar surface are desired. At a slant range of 60 miles, the altitude marking radar generates a signal that, after a suitable time delay established by ground command, initiates the ignition of the vernier engines that are used to control the spacecraft attitude during the main retro burning. One second after vernier ignition, the solid main retro motor is ignited. After 41 seconds thrust decay is sensed by an acceleration switch that initiates a time delay of 12 seconds and the empty main retro case is jettisoned.

Shortly after separation, vernier engine thrust is reduced to 0.9 lunar g’s. The radar altimeter and doppler velocity sensor (RADVS) senses the lunar surface and provides measurements of range and velocity to the flight control system, and the thrust axis is aligned with the velocity vector. When the RADVS senses that the pre-programmed range-velocity contour has been reached, the thrust is increased to an acceleration level such that the spacecraft descends along this contour until a speed of 10 fps is reached. The dynamic character of the descent trajectory (a “gravity turn trajectory”) is such that when the spacecraft reaches the 10 fps point, the thrust attitude will be nearly aligned to the vertical. The vehicle attitude is then inertially fixed and the maximum acceleration is held down to an altitude of 13 feet when the engines are shut off and the vehicle free falls to the surface with a landing speed of 13 feet/second.

Surveyor Terminal Descent and Soft Landing

After landing on the surface of the moon, the spacecraft is commanded to align the solar panel to the sun-line and the high grain antenna to earth. Subsequent commands connect one of the transmitters to the to the planar array antenna and switch that mode (high or low power) that will provide the necessary communications bandwidth consistent with the operation being performed, the power available fro the solar panel and battery, and the thermal restrictions imposed by the compartment dissipating capabilities. At this point, any of various engineering or scientific experiment sequences may commence.

SCG in line to build GOES satellite—Hughes News August 19, 1977 transcribed by Faith Macpherson

Space and Communications Group soon will begin negotiations with the National Aeronautics and Space Administration (NASA) to build, test and deliver three Geostationary Operational Environmental Satellites (GOES).

GOES D, E, and F, each designed for a seven year orbital life, are continuations of the Synchronous Meteorological Satellite-GOES series.

When complete, the system will provide near continuous, timely, and high quality observations of the Earth and its environment.

The eventual contract for the three spacecraft is expected to be worth approximately $33 million.

Scheduled for launch in mid-1980, GOES D is planned to be one of the first operational payloads to be launched from NASA’s Space Shuttle Orbiter. The satellite is being designed for launch from either the Orbiter or a Delta class rocket.

“We are very pleased with the award,” said Harvey Palmer, NASA Systems Division manager.

“GOES D, E and F will compliment out already existing Geosynchronous Meteorological Satellite (GMS) series which we are developing for the Japanese.”

GMS I was successfully launched on July 14 and negotiations are expected to begin soon for GMS II.

“GOES D will be the first spacecraft to make use of a Visible Infrared Spin Scan Radiometer Atmospheric Sounder (VAS),” said Steve Petrucci, proposal manager.

“The camera, under production at Santa Barbara Research Center, will provide data indicating the temperature versus altitude profile of the atmosphere, as well as provide pictures of the visible cloud cover and the infrared capabilities of earlier cameras.”

The payload will contain a Space Environmental Monitoring System consisting of three separate sensors designed to monitor solar emission activities.

Each drum-shaped spacecraft will be 85 inches in diameter at the widest point and 175 inches in length.

SCG hosts Jupiter probe progress review—Hughes News February 2, 1979 transcribed by Faith Macpherson

HAC’s Probe spacecraft for NASA’s Galileo mission to Jupiter was the subject of a three-day design review hosted by Space and Communications Group’s NASA Systems Division.

The Probe is one of two Galileo mission spacecraft due to be launched from the Space Shuttle in January 1982 on a 3 ½-year, 500-million mile journey to the largest planet in the solar system. The mission is intended to provide scientists with detailed information about Jupiter and will provide the first close-up view of the planet’s atmosphere.

NASA Systems Division, under a $35 million contract with NASA, is developing the Probe and the telemetry relay communications receiver and antenna that will be aboard the second spacecraft, the Orbiter.

NASA Ames Research Center is managing the Galileo Program and the Jet Propulsion Laboratory has overall project management responsibility and is developing the Orbiter.

Representatives from NASA Headquarters in Washington, DC, NASA Ames, Hughes, JPL, and General Electric attended the review.

The Probe, which is made up of the descent module and the deceleration module, weighs 550 pounds, is 49 inches in diameter, and is 34 inches long.

SCG is developing the descent module of the Probe and has subcontracted the development of the deceleration module to General Electric’s Re-Entry Systems Division in Philadelphia. The deceleration module is a heat shield that will protect the descent module until it enters the planet’s atmosphere.

The descent module is designed to descend through Jupiter’s atmosphere to the planet, transmitting the findings of the scientific instruments it carries to the Orbiter, which will relay the information to Earth.

The Orbiter will circle the planet for 20 months, taking pictures and analyzing the atmosphere from its vantage point 622,000 miles away from the planet.

The descent module will carry all the electronics and scientific instruments for the mission.

The Probe will separate from the Orbiter 150 days away from Jupiter. Instruments will be turned on five hours before the Probe reaches the planet.

The descent module will ride inside the deceleration module during those 150 days until they reach Jupiter’s atmosphere. The descent module will then be pulled out of its protective “capsule” by a parachute. The parachute will help slow the descent module’s pass through the atmosphere, which will allow for a longer sampling time, according to Uldis Lapins, HAC Program manager.

Radio transmissions back to Earth from the Galileo spacecraft will begin once the probe has entered the atmosphere.     Six instruments will be aboard the descent module. The instruments and their objectives are:

• an atmosphere structure instrument to chart the probe’s deceleration history and then give temperature and pressure profiles of the atmosphere.
• a nephelometer will show the vertical extent and structure of the clouds.
• a helium abundance detector will show the ratio of helium to hydrogen in the atmosphere.
• a net flux radiometer will measure the planet’s energy balance, which is the relationship between energy from the Sun reaching the planet and the energy coming from the planet.
• a lightning and radio emission detector will measure the amount of radio frequency noise.
• a neutral mass spectrometer will chart the atmosphere’s chemical compounds and physical state.

On their way to Jupiter, the Galileo spacecraft will come within a few thousand kilometers of Mars.