Instrument Host Overview - Spacecraft
=====================================

Launch Vehicle Description
--------------------------

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.

Flight System Description
-------------------------

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.

Mechanical Structure
--------------------

The Mars Pathfinder flight system was effectively three spacecraft in
one.  The flight system was a standard interplanetary spacecraft in
one mode, an atmospheric entry vehicle in another, and a surface
lander in the third.  The major components of the spacecraft were the
cruise stage, the entry vehicle (consisting of the heatshield,
backshell, and attached hardware), and the lander (containing the
rover).  Components from all other subsystems were distributed through
each of these elements.

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.

Entry, Descent, and Landing Subsystem
-------------------------------------

The Entry, Descent, and Landing subsystem was responsible for safely
placing the lander on the surface of Mars.  In order to do this, the
velocity had to be reduced from the initial entry velocity of 7.6 km/s
to a level that limited the maximum landing shock to less than 40 g's.
Specific components of the system required to do this included a 2.65
meter diameter aeroshell, accelerometers, a parachute, an incremental
bridle, retrorockets, airbags, and a radar altimeter.

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/m^2.

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.


Attitude and Information Management Subsystem
---------------------------------------------

The AIM subsystem performed all spacecraft computing functions
including command and telemetry handling, HGA pointing, payload data
compression, cruise attitude determination and control, and EDL
timing.  This subsystem was built around a single high-performance (20
MIPS), 32-bit, single-board R6000 computer.  The computer had 128
Mbyte of dynamic RAM for flight software, engineering measurements,
and science data.  An additional 4 Mbytes of electronically erasable
PROM was included to hold flight software boot code, critical
sequences and data.  The speed and capabilities of this computer
greatly simplified both software development and mission operations.

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.

Telecommunications Subsystem
----------------------------

The Telecommunications subsystem provided communications capability
during cruise, EDL and surface operations and support for radio
navigation during cruise.  The baseline design was a direct to Earth
X-band system.  There were two major elements to this subsystem: the
Radio Frequency Subsystem (RFS) and the Antenna Subsystem.  All RFS
components were located in the lander ISA.  Primary RFS functions were
performed by a single string Cassini transponder, a 13 watt (RF) Solid
State Power Amplifier, a newly developed Telemetry Modulation Unit,
and a Cassini Command Decoder Unit.  This single string approach was
designed to work for the duration of the short Pathfinder mission,
however, a partially redundant backup downlink capability was also
provided.  This backup system included a small X-band transmitter with
an integrated 5 W (RF) power amplifier and an additional TMU.  This
backup provided significantly less performance than the primary
system, but was designed to be sufficient to satisfy the primary
mission objectives in an emergency.

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.

Power and Pyrotechnics Subsystem
--------------------------------

The Power and Pyrotechnics Subsystem was responsible for generating,
storing and distributing power during cruise and surface operations.
In addition, this subsystem was responsible for controlling and
initiating all pyrotechnic devices required for EDL and surface
operations.  Power was generated during cruise by a 4.4 m^2 Gallium
Arsenide (GaAs) solar array which covered the top surface of the
cruise stage.  The 5.5 mm thick cells were arranged to prevent any
single failures from catastrophically impacting the power output.
Power was generated on the surface by 2.9 m^2 of GaAs solar arrays
mounted on the exposed surfaces of the lander petals.  The lander and
cruise stage used the same 5.5 mm thick cells.

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.

Propulsion Subsystem
--------------------

The Propulsion subsystem included all equipment needed to perform
attitude control and TCM's during cruise and Rocket Assisted
Deceleration during EDL.  The cruise stage propulsion system consisted
of four blow-down hydrazine titanium propellant tanks connected with
eight 4.45 N (1 lb) thrusters via series latch valves.  These
thrusters were arranged to allow coupled turns and both axial and
transverse translational maneuvers.  The purpose and function of the
RAD rockets are described more fully in the EDL section above.

Mass Summary
------------

The spacecraft launch mass was 894 kg, including the hydrazine (N2H4)
propellant, science instruments, and a free-ranging rover surface
vehicle. The required amount of propellant was calculated using the
maximum available spacecraft launch mass of 905 kg. The entry mass was
584 kg. The entry mass was a key driver on the EDL system design, so
careful margin management was important.

Spacecraft Summary
------------------

Launch Mass:        894 kg (Includes Propellant)
Entry Mass:         584 kg
Lander Mass:        370 kg

Basic Design:
   - Aeroshell, parachute, RAD rocket and airbag Entry, Descent,
     and Landing (EDL) system
   - Self righting, tetrahedral lander
   - Active thermal design for cruise
   - Free ranging rover

Command And Data Handling
   - Integrated Attitude and Information Management System (AIM)

Computation
   - R6000 Computer with VME bus, 22 Millions of Instructions Per
     Second (MIPS), 128 Mbyte mass memory

Power
   - Solar powered cruise stage and lander

Telemetry And Command
   - Surface operations telemetry rate via High Gain Antenna (HGA),
     X-Band: 6 kb/s to 70 m Deep Space Network
   - Surface operations command rate via HGA, X-Band: 250 b/s

Propulsion
   - Monopropellant hydrazine used for cruise
   - Eight 4.4 N thrusters
   - Total delta-v of 130 m/s

Instrument Host Overview - DSN
==============================

The Pathfinder radio telecommunications system was used to monitor the
distance between Earth and Pathfinder and its rate of change in order
to determine the rate of precession of Mars.  This utilized elements
of the NASA Deep Space Network (DSN) in addition to the
instrumentation on board the spacecraft.

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].