The launch vehicle used for Mars Pathfinder was the McDonnell Douglas
Delta II 7925. An engine section in the Delta first stage housed the
Rocketdyne RS-27 main engine and two Rocketdyne LR101-NA-11 vernier
engines. The vernier engines provided roll control during main engine
burn, and attitude control after main engine cutoff and before second
stage separation. The RS-27 main engine was a single start, liquid
bipropellant rocket engine which provided approximately 894,094 N of
thrust at liftoff. The first stage propellant load (96,000 kg)
consisted of RP-1 fuel (thermally stable kerosene) and liquid oxygen
as an oxidizer. The RP-1 fuel tank and liquid oxygen tank were
separated by a center body section that housed control electronics,
ordnance sequencing equipment, a telemetry system, and a rate gyro.
First stage thrust augmentation was provided by nine solid-propellant
Graphite Epoxy Motors (GEMs), each fueled with 12,000 kg of
hydroxyl-terminated polybutadiene solid propellant. Each GEM provided
an average thrust of 439,796 N at liftoff. The main engine, vernier
engines, and six of the GEMs were ignited at liftoff. The remaining
three GEMs were ignited in flight. The GEMs were jettisoned from the
first stage after motor burnout. The interstage assembly extended from the top of the first stage to the second stage mini-skirt. This assembly carried loads to the first stage, and contained an exhaust vent and six spring driven separation rods.
The Delta II 7925 second stage propulsion system included a restartable, liquid bipropellant Aerojet AJ10-118K engine that consumed Aerozine 50 fuel (a 50/50 mix of hydrazine and unsymmetric dimethyl hydrazine) and nitrogen tetroxide (N2O4). Since Aerozine 50 and nitrogen tetroxide are hypergolic, no catalyst or igniter in the engine thrust chamber was required. The second stage had a total propellant load of 6000 kg and provided a thrust of 42,000 N. Gaseous helium was used for pressurization, and a nitrogen cold gas jet system provided attitude control during coast periods and roll control during powered flight. Hydraulically activated gimbals provided pitch and yaw control. The forward section of the second stage housed the guidance and control equipment that provided guidance sequencing and stabilization signals for both the first and second stages. Both first and second stages had a battery driven DC power system. Separate batteries were used for the guidance and control system, ordnance, and engine systems. The instrumentation and flight termination system were powered by the same battery.
The PAM-D was the third stage of the Delta II 7925 launch vehicle and provided the final velocity required to place the Pathfinder spacecraft on a trajectory to Mars. The PAM-D upper stage consisted of (1) a spin table to support, rotate, and stabilize the Pathfinder spacecraft / PAM-D combination before separating from the second stage, (2) a Star-48B solid rocket motor for propulsion, (3) an active Nutation Control System (NCS) to provide stability after spin-up of the spacecraft/PAM-D stack, (4) a 3717C payload attach fitting to mount the Star-48B motor to the spacecraft, and (5) a yo-yo de-spin system designed to decrease the spin rate of the upper stage / spacecraft stack from 70 RPM to 12 ± 2 RPM. The Star-48B provided an average thrust of 67,168 N and was fueled with 2000 kg of solid propellant that was composed primarily of ammonium perchlorate and aluminum.
During launch and ascent through the sensible atmosphere, the Pathfinder spacecraft / PAM-D upper stage combination was protected from aerodynamic forces by a 2.9 m diameter payload fairing. The payload fairing was jettisoned from the launch vehicle during second stage powered flight at a minimum altitude of 111 km.
The Mars Pathfinder Flight System consisted of six major subsystems:
Attitude and Information Management (AIM); Telecommunications; Entry,
Descent, and Landing (EDL); Power and Pyrotechnics; Propulsion; and
Mechanical Structure. Brief descriptions of the subsystems are given
below.
The purpose of the lander was to provide engineering support to the
science instruments and rover during surface operations. The landing
approach employed by Mars Pathfinder required the lander to have self
righting capability. A simple tetrahedron design was developed to
limit the possible landing orientations. The lander was able to right
itself if it landed on any face of the tetrahedron by driving
actuators which connected three faces of the tetrahedron (the petals)
with the fourth (the base petal). All thermally sensitive electronics
were contained in the insulated Integrated Support Assembly (ISA)
located inside the tetrahedron. This enclosure provided a controlled
environment to minimize the effects of the extreme temperature
variations on Mars. The lander high gain antenna and low gain antenna
were attached to the outside of the ISA, as was the Imager for Mars
Pathfinder. The rover was attached to one of the side petals.
A detailed description of the entry vehicle is given in the following EDL subsystem description. The cruise stage was responsible for controlling the vehicle during cruise, and was home to most of the components required for interplanetary flight. These included launch vehicle separation devices; attitude sensors; propulsive thrusters, tanks, and other equipment; cruise antennas; and solar arrays. The cruise stage was jettisoned prior to entry into the Martian atmosphere and impacted the surface. To eliminate duplication of capabilities, cruise stage hardware utilized key telecommunications, attitude and information management services, and power support from equipment that resided on the lander. As a result, there was a significant umbilical extending between the lander and cruise stage.
The aeroshell was used to reduce the velocity of the vehicle from 7.6 km/s to .370 km/s, removing over 99% of the initial kinetic energy and protecting the lander against the resultant extreme aerodynamic heating. The heatshield was a Viking-heritage, 70 degree half angle blunt cone. Thermal protection was provided by Viking-heritage SLA-561V ablative material. Ablative material was also applied to the backshell to protect the lander from the effects of recirculation flow around the entry vehicle. During the entry and descent phases, accelerometers on the lander provided parachute deployment timing information and acceleration data. A 12.7 meter diameter Viking heritage disk-gap-band parachute was used for terminal descent. The parachute was designed to open at supersonic speeds with a maximum deployment dynamic pressure of 700 N/m2.
The EDL system also included separation devices designed to separate the heatshield from the backshell after parachute deployment. The heatshield had to be separated and released from the backshell so that the lander could be deployed on the incremental bridle. The incremental bridle was designed to provide separation between the backshell-mounted retrorockets and the lander and to improve stability during the rocket firing. A radar altimeter on the lander was used to determine when to ignite the three retrorockets. The retrorockets were sized to bring the backshell/lander system to a complete stop at approximately 15 m above the surface. Four airbags attached to the faces of the tetrahedral lander were inflated just before the rockets fired and were designed to limit landing loads to less than 40 Earth g's. The airbags were inflated using three gas generators in 1/4 of a second. The generators continued to maintain the pressure in the airbags for over one minute past initial pressurization. Just prior to contacting the surface, cable cutters released the lander from the parachute, backshell, and retrorockets. The lander hit with a vertical velocity of 12 m/s and a horizontal velocity of 6 m/s. After ground impact and tumbling, the airbags deflated and were retracted. The three lander petals were then opened to establish an upright configuration on the surface.
All of the details of the timing for the entire EDL procedure were determined in real time on board the spacecraft.
Internal interfaces within AIM were implemented using a VME bus backplane. Major VME interfaces included the Power & Pyrotechnics subsystem, the Reed Solomon downlink board, the hardware command decoder, and the power converter unit. A Mil-Std-1553 bus served for engineering data links to remote engineering units which collected status data from the celestial sensors and analog telemetry channels. Celestial sensors included a modified Magellan/IUS star scanner and a five head sun sensor. The spacecraft was spin stabilized during cruise, and maintained an Earth point profile for most of the flight (except for portions near Earth and Mars). The attitude determination and control system was required to maintain at least 2 degrees inertial pointing throughout cruise. Attitude determination and HGA control was required during the surface phase. The autonomous pointing accuracy requirement for the HGA was 3.5 degrees (note that more accurate pointing was possible using downlink dithering). The AIM was also responsible for collecting and packetizing science data from the instruments.
Flight software in the AIM controlled the uplink interface, downlink telemetry, instrument control, engineering / science data collection and formatting, bus control, sequence and command processing, attitude determination / control, and power / pyro functions. The flight software design was coded in 'C'. The operating system was an adaptation of VxWorks, a commercially-available operating system for the R6000 computer. The operating system provided a file system and an interprocess communications protocol for flight software components.
The Antenna Subsystem consisted of five antennas for cruise, EDL and surface operations. A waveguide connected the RFS in the lander to a medium gain antenna (MGA) located on the cruise stage. A medium gain antenna was required to maintain a minimum 20 b/s link through cruise. The MGA was a standard Mars Observer design. Two antennas were provided for EDL communications. The backshell LGA was used during entry and early parachute descent. The EDL LGA was a whip antenna attached to the top of the lander and was used during the final portion of parachute descent. Two antennas were provided on the lander. The principal antenna was a steer-able high gain antenna mounted on the ISA. The HGA was a mechanically-steered slotted plate with 2 degrees of freedom in pointing. It provided a nominal 125-b/s command uplink rate and a telemetry downlink rate of approximately 600 b/s into a DSN 34-m antenna (or 2700 b/s into a 70-m antenna). These data rate capabilities assumed 3.5 degrees pointing of the HGA. Improved pointing (using a dithering scheme with the DSN) could improve these rates. These rates were also conservative in that they contained a 3 dB link margin. A Low Gain Antenna was provided for emergency situations in the event the HGA should fail. The LGA provided an emergency 7.8 b/s command uplink rate and a minimum 40 b/s downlink rate over a 60 degree beamwidth.
A 40 amp-hr (~1120 W-hr @ 28 V) silver zinc battery was used to provide energy storage during the mission. Silver zinc batteries are typically used in primary battery applications (like launch vehicles), but Pathfinder re-qualified the technology for use as a rechargeable battery. The advantage of this type of battery is a much higher energy density than typical NiCd or NiH batteries. The disadvantage is a relatively short cycle life (limited to 30-100 cycles). The battery was only used to support launch, trajectory correction maneuvers (TCMs) and surface operations, however, so the cycle capability was sufficient for primary operations.
The power distribution system used an unregulated bus controlled by a shunt regulator. Pyro devices were initiated using a secondary power distribution system (isolated from the main) to prevent ground loop problems. Because of the rapid firing of pyros during EDL, special thermal batteries were included as a backup power source for the pyro system. The pyro system was designed to provide sufficient arm and enable controls to satisfy all system safety requirements.
| Launch Mass: | 894 kg (Includes Propellant) |
|---|---|
| Entry Mass: | 584 kg |
| Lander Mass: | 370 kg |
Basic Design:
Command And Data Handling:
Computation:
Power:
Telemetry And Command:
Propulsion:
The Deep Space Network is a telecommunications facility managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration.
The primary function of the DSN is to provide two-way communications
between the Earth and spacecraft exploring the solar system. To carry
out this function the DSN is equipped with high-power transmitters,
low-noise amplifiers and receivers, and appropriate monitoring and
control systems.
The DSN consists of three complexes situated at approximately equally spaced longitudinal intervals around the globe at Goldstone (near Barstow, California), Robledo (near Madrid, Spain), and Tidbinbilla (near Canberra, Australia). Two of the complexes are located in the northern hemisphere while the third is in the southern hemisphere.
The network comprises four subnets, each of which includes one antenna at each complex. The four subnets are defined according to the properties of their respective antennas: 70-m diameter, standard 34-m diameter, high-efficiency 34-m diameter, and 26-m diameter.
These DSN complexes, in conjunction with telecommunications subsystems onboard planetary spacecraft, constitute the major elements of instrumentation for radio science investigations.
For more information see [ASMAR&RENZETTI1993].
More information about the Deep Space Network is available at the following WWW site: http://deepspace1.jpl.nasa.gov/dsn/.