Virginia Polytechnic Institute and State University Venus Sample Return Space Design Team

Virginia Polytechnic Institute and State University
Venus Sample Return
Space Design Team
May 4, 2001
For Submission to:
2000/2001 AIAA Foundation Undergraduate Space Design Competition
Dr. Christopher Hall
Team Members:
Giuseppe Angelone
Victor Collazo-Perez
Greg Evans
Gregory Fertig
Kathleen Hale
Laird-Philip Lewis
Amy Spratley
James St. Amand
Andrew Wallace
Table of Contents
List of Figures.................................................................................................................... iv
List of Tables ..................................................................................................................... vi
List of symbols ................................................................................................................. vii
Chapter 1 - Introduction.......................................................................................................1
1.1 - Mission Summary .............................................................................................................................. 1
1.2 - Venus Science Information................................................................................................................ 1
Chapter 2 - Mission Concepts..............................................................................................3
2.1 - Constrains .......................................................................................................................................... 3
2.2 - Propulsion Systems............................................................................................................................ 3
2.2.1 - Orbiter ........................................................................................................................................ 3
2.2.2 - Venus Insertion Package ............................................................................................................ 3
2.2.3 - Venus Ascent Vehicle ................................................................................................................ 3
2.2.4 - Earth Insertion Package.............................................................................................................. 4
2.3 - Entry Systems .................................................................................................................................... 4
2.3.1 - Venus Entry................................................................................................................................ 4
2.3.2 - Earth Entry ................................................................................................................................. 4
2.4 - Attitude Determination and Control Systems .................................................................................... 4
2.5 - Thermal.............................................................................................................................................. 5
2.5.1 - Orbiter ........................................................................................................................................ 5
2.5.2 - Venus Lander ............................................................................................................................. 5
2.6 - Mechanisms ....................................................................................................................................... 5
2.6.1 - Orbiter and Earth Entry Vehicle................................................................................................. 5
2.6.2 - Venus Lander ............................................................................................................................. 6
2.7 - Computer / Communications ............................................................................................................. 7
2.7.1 - Computer.................................................................................................................................... 7
2.7.2 - Communications ........................................................................................................................ 7
2.8 - Rendezvous........................................................................................................................................ 7
2.9 - Power ................................................................................................................................................. 8
Chapter 3 - Main Orbiter Bus ............................................................................................11
3.1 - Configuration ................................................................................................................................... 11
3.1.1 - Heliogyro and Support Structure.............................................................................................. 11
3.1.2 - Main Bus .................................................................................................................................. 13
3.1.3 - Aeroshell .................................................................................................................................. 14
3.2 - Thermal............................................................................................................................................ 15
3.3 - Attitude Determination and Control System.................................................................................... 17
3.3.1 - Attitude Determination............................................................................................................. 17
3.3.2 - Control Systems ....................................................................................................................... 17
3.4 - Power ............................................................................................................................................... 22
3.5 - Computer / Communication............................................................................................................. 23
3.5.1 - Computer.................................................................................................................................. 23
3.5.2 - Communications ...................................................................................................................... 24
3.6 - Propulsion ........................................................................................................................................ 24
3.6.1 - Solar Sailing Basics.................................................................................................................. 25
3.6.2 - Equations of Motion................................................................................................................. 26
3.6.3 - Interplanetary Travel ................................................................................................................ 28
3.6.4 - Travel Around Venus ............................................................................................................... 30
3.6.5 - Future Analysis ........................................................................................................................ 32
3.7 - Mechanisms ..................................................................................................................................... 33
3.7.1 - Lightband ................................................................................................................................. 33
3.7.2 - Solar Sail Blade Thrusters........................................................................................................ 33
3.7.3 - Communications Dish Pointing Mechanism ............................................................................ 33
3.7.4 - Blade Rotation Motors ............................................................................................................. 33
Chapter 4 - Venus Lander..................................................................................................35
ii
4.1 - Configuration ................................................................................................................................... 35
4.2 - Sizing Methodology......................................................................................................................... 36
4.2.1 - Helium tanks ............................................................................................................................ 36
4.2.2 - Titanium Platform .................................................................................................................... 38
4.2.3 - Landing Legs............................................................................................................................ 38
4.2.4 - Center of Mass ......................................................................................................................... 39
4.3 - Thermal............................................................................................................................................ 40
4.4 - Attitude Determination and Control Systems .................................................................................. 42
4.5 - Power ............................................................................................................................................... 43
4.6 - Computer ......................................................................................................................................... 44
4.6.1 - Venus Lander Computer .......................................................................................................... 44
4.6.2 - Sample Capsule Computer ....................................................................................................... 44
4.7 - Mechanisms ..................................................................................................................................... 44
4.7.1 - Ultrasonic Drill/Corer .............................................................................................................. 44
4.7.2 - Mechanical Arm and Scoop ..................................................................................................... 45
4.7.3 - Sample Containers ................................................................................................................... 46
4.8 - Scientific Instrumentation ................................................................................................................ 47
4.8.1 - Variometer ............................................................................................................................... 47
4.8.2 - Wind Vane ............................................................................................................................... 47
4.8.3 - Panoramic Micro-Imager ......................................................................................................... 47
4.9 - Venus Entry and Descent................................................................................................................. 47
4.9.1 - Ballute Introduction ................................................................................................................. 47
4.9.2 - Shape........................................................................................................................................ 48
4.9.3 - Materials................................................................................................................................... 49
4.9.4 - Sizing ....................................................................................................................................... 51
4.9.5 - Trajectory ................................................................................................................................. 52
4.9.6 - Post-Entry Descent................................................................................................................... 55
4.10 - Venus Ascent ................................................................................................................................. 57
4.10.1 - Venus Ascent Vehicle (Balloon)............................................................................................ 57
4.10.1.a - Material Selection........................................................................................................... 57
4.10.1.b - Shape and Size................................................................................................................ 60
4.10.1.c - Balloon Ascent ............................................................................................................... 62
4.10.2 - Venus Ascent Vehicle (Rocket) ............................................................................................. 64
Chapter 5 - Earth Entry Vehicle ........................................................................................72
5.1 - Configuration ................................................................................................................................... 72
5.1.1 - Sample Collector ...................................................................................................................... 72
5.2 - Thermal............................................................................................................................................ 73
5.3 - Attitude Determination and Control Systems .................................................................................. 73
5.4 - Power ............................................................................................................................................... 74
5.5 - Computer ......................................................................................................................................... 74
5.6 - Propulsion ........................................................................................................................................ 74
5.7 - Earth Entry and Descent .................................................................................................................. 75
5.8 - Sample Analysis .............................................................................................................................. 76
Chapter 6 - Cost Analysis ..................................................................................................77
References..........................................................................................................................79
Appendix A – Mission Timeline .............................................................................................................. 83
Appendix B – Venus Lander Schematic .................................................................................................. 84
Appendix C – Orbiter Schematic ............................................................................................................. 85
iii
List of Figures
Figure 1 – Blade Taper ................................................................................................................................. 11
Figure 2 - Stowed Orbiter Configuration ...................................................................................................... 11
Figure 3 - Hexagonal Ring Structure ............................................................................................................ 12
Figure 4 - Blade Arm Deployment Procedure .............................................................................................. 12
Figure 5 - Intermediate Orbiter Deployment................................................................................................. 13
Figure 6 - Deployed Orbiter.......................................................................................................................... 13
Figure 7 - Venus Lander Aeroshell............................................................................................................... 14
Figure 8 – Hypersonic Shock Wave Profile.................................................................................................. 15
Figure 9 - Hypersonic Shock Wave Through Ballute ................................................................................... 15
Figure 10 - Multi-Layered Insulation Cross-Section .................................................................................... 16
Figure 11 - Radial (Lengthwise) Stress Analysis.......................................................................................... 19
Figure 12 - Tensile Stress Along Blade Chord ............................................................................................. 19
Figure 13 - Coning Angle versus Position Along Length of Blade .............................................................. 20
Figure 14 - Blade Shape versus Position Along Length of Blade ................................................................. 21
Figure 15 - Required Torque versus Pitch Angle.......................................................................................... 22
Figure 16 - RHPPC Mechanical Concept (Ref #9) ....................................................................................... 24
Figure 17 - System before photon strike ....................................................................................................... 25
Figure 18 - System after photon strike.......................................................................................................... 25
Figure 19 - Polar Coordinates Defined ......................................................................................................... 26
Figure 20 - Mean thrust for travel to Venus.................................................................................................. 28
Figure 21 - Travel Trajectory From Earth to Venus at Minimum Travel Time Conditions.......................... 29
Figure 22 - Travel Trajectory From Venus to Earth at Minimum Travel Time Conditions.......................... 29
Figure 23 - Overview of Venus Capture ....................................................................................................... 30
Figure 24 - Venus Capture Close-up............................................................................................................. 31
Figure 25 - Venus Escape Trajectory............................................................................................................ 32
Figure 26 - Deployed Venus Lander ............................................................................................................. 35
Figure 27 - Stowed Venus Lander ................................................................................................................ 36
Figure 28 - Shock Absorber Deployed and Stowed Configurations ............................................................. 36
Figure 29 - Venus Lander Main Platform ..................................................................................................... 38
Figure 30 - Venus Lander Leg Deployed Configuration .............................................................................. 39
Figure 31 - Venus Lander Center of Mass Layout........................................................................................ 40
Figure 32 - Venus Lander Thermal Shields .................................................................................................. 40
Figure 33 - Venus Thermal Shielding........................................................................................................... 41
Figure 34 - Venus Shielding Heat Transfer versus Time .............................................................................. 42
Figure 35 - Thermal Conductivity versus Temperature ................................................................................ 42
Figure 36 - Close up of the Ultrasonic Drill/Corer ....................................................................................... 45
Figure 37 - Attached Aeroshell (Ref #18)..................................................................................................... 48
Figure 38 - Torroidal Ballute and Aeroshell ................................................................................................. 49
Figure 39 - Cross Section of Torroidal Ballute ............................................................................................. 49
Figure 40 - Ballute with Final Dimensions ................................................................................................... 51
Figure 41 - Venus Entry Trajectory .............................................................................................................. 53
Figure 42 - Entry Sensitivity......................................................................................................................... 54
Figure 43 - Entry Deceleration and Density vs. Altitude .............................................................................. 54
Figure 44 - Velocity and Density verersus. Altitude..................................................................................... 55
Figure 45 - Descent Altitude vs. Time .......................................................................................................... 56
Figure 46 - Descent Velocity vs. Time ......................................................................................................... 56
Figure 47 - Chemical structure of PBO (Ref #42) ........................................................................................ 58
Figure 48 - Strength and Modulus vs. Temperature (Smith) ........................................................................ 58
Figure 49 - Helium Permeability of Several Possible Balloon Materials...................................................... 59
Figure 50 - Balloon Seam (from 99-3858).................................................................................................... 60
Figure 51 - Balloon with both payload attachments ..................................................................................... 62
Figure 52 - Lifting Gas Analysis................................................................................................................... 63
iv
Figure 53 - Ascent Altitude vs. Time............................................................................................................ 64
Figure 54 - Venus Ascent Vehicle Concept .................................................................................................. 67
Figure 55 - Venus Ascent Vehicle Dimensions ............................................................................................ 68
Figure 56 - Venus Ascent Vehicle Flight Path Profile.................................................................................. 69
Figure 57 - Venus Ascent Vehicle Launch Profile ....................................................................................... 70
Figure 58 - Venus Ascent Vehicle Altitude versus Time Plot ...................................................................... 71
Figure 59 - Orbiter, EEV, with Extended Cone ............................................................................................ 72
Figure 60 - DSBC Computer ........................................................................................................................ 74
Figure 61 - Orbiter and Earth Entry Vehicle Separation............................................................................... 75
Figure 62 - Earth Entry Vehicle with Descent Parachutes ............................................................................ 76
v
List of Tables
Table 1 - Attitude Determination and Control System Summary ................................................................... 5
Table 2 - Orbiter Mechanisms ........................................................................................................................ 6
Table 3 - Lander Instruments .......................................................................................................................... 6
Table 4 - Power Requirements for Spacecraft Components ........................................................................... 9
Table 5 - Summary of Power Sources........................................................................................................... 10
Table 6 - Temperature Ranges for Sensitive Components (Ref #39) ........................................................... 16
Table 7 - Orbiter Batteries ............................................................................................................................ 22
Table 8 - RHPPC Feature Summary (Ref #9)............................................................................................... 23
Table 9 - Radiation Hardness (Ref #9) ......................................................................................................... 24
Table 10 - Helium Tank Geometry and Mass Combinations........................................................................ 37
Table 11 - Venus Lander Batteries (Ref #32) ............................................................................................... 43
Table 12 - Ballute Film Materials (Ref #42)................................................................................................. 50
Table 13 - Ballute Fiber Materials (Ref #42)................................................................................................ 50
Table 14 - Tensile Stress Analysis of Kapton and PBO (Ref #22) ............................................................... 50
Table 15 - Final Ballute Materials and Masses ............................................................................................. 52
Table 16 - Balloon material comparison (Ref #35)...................................................................................... 57
Table 17 - Possible Corrosive Protection Materials..................................................................................... 60
Table 18 - Initial Balloon Sizing Analysis .................................................................................................... 61
Table 19 - Final Balloon Specifications........................................................................................................ 63
Table 20 - Propellant Performance Characteristics (Ref #17 p353).............................................................. 65
Table 21 - Material Properties of Graphite (Ref #17 p310) .......................................................................... 65
Table 22 - Venus Ascent Vehicle Stage One Configuration......................................................................... 66
Table 23 - Venus Ascent Vehicle Stage Two Configuration ........................................................................ 67
Table 24 - Apparent Thermal Conductivity (Ref #41).................................................................................. 73
Table 25 - Performance Characteristics of Propulsion Systems (Ref #19 p692) .......................................... 75
Table 26 - Component Costs......................................................................................................................... 78
Table 27 - Fabrication Costs ......................................................................................................................... 78
vi
List of symbols
Variable
As
c
CD
cg
cpa
cps
∆Τ
∆V
Fs
g
g0
γ
h
Isp
µ
P
p
q
r
R
Rearth
Rvenus
q
θA
t
Tsp
V
Ve
Description
Cross-sectional Area
Speed of Light (3 × 108 m/s)
Drag Coefficient
Center of Gravity
Aerodynamic Center
Center of Solar Pressure
Total transfer time
Change in Velocity
Solar constant (1,367 W/m2)
Acceleration due to gravity
Acceleration due to gravity on Earth’s surface (9.8 m/s2)
Flight path angle
Angular Momentum
Specific Impulse
Gravitational constant
Orbital Period
Orbital Parameter
Reflectance Factor
Atmospheric Density
Orbital Radius
Radius of Earth
Radius of Venus
Maximum Deviation of Z-axis from Local Vertical
Allowable Motion
Time
Maximum Solar Radiation Pressure Torque
Velocity
Exit velocity
vii
Chapter 1 - Introduction
1.1 - Mission Summary
The mission to Venus involves a complex set of equipment and maneuvers. A Delta IV Medium Plus (5m)
lifter rocket is used to send the Venus spacecraft out of the Earth’s influence. A heliogyro solar sail
transports the spacecraft to Venus. The twelve solar sail blades deploy once the Venus spacecraft is
separated from the Delta IV upper stage. This heliogyro device is a propulsion system that allows the mass
of the spacecraft to be significantly lower than a craft using chemical propulsion. The heliogyro is used to
maneuver the spacecraft to rendezvous with Venus 452 days after leaving Earth.
The orbiter maneuvers into a Venus orbit with an 800-km periapsis and 275,000 km apoapsis. Once the
orbiter reaches this orbit the lander, inside its aeroshell, is detached and sent into Venus’s atmosphere. The
lander enters Venus’s atmosphere and deploys a ballute to slow its descent. Once the lander reaches an
altitude of about 70-km the ballute detaches with the upper aeroshell and releases a balloon. The balloon is
used to slow the lander as it descends to the surface.
During descent, atmospheric samples are taken and wind direction consistency is recorded. An ultrasonic
corer and mechanical arm are used to acquire a two-kilogram surface sample once the lander reaches the
surface. The balloon then lifts a rocket containing the sample to an altitude of about 61-km. The rocket
launches and transports the sample into an 800-km orbit. The orbiter collects the sample and the spacecraft
returns to Earth.
The travel time from Venus to Earth is approximately 119 days, once again using the heliogyro solar sail for
propulsion. The Earth Entry Vehicle (EEV) is detached and sent into Earth’s atmosphere along with the
sample and collected data. The sample lands in the Pacific Ocean and is retrieved for analysis.
See Appendix A for timeline.
1.2 - Venus Science Information
The most challenging obstacle to overcome in a Venus surface mission is preparing for the planet’s
environmental conditions. The Venusian environment is among the harshest in the solar system. The
atmosphere is 96% carbon dioxide, 3.5% nitrogen, and 0.5% trace compounds, including carbon monoxide,
sulfuric acid, hydrochloric acid, and hydrofluoric acid. The high amount of carbon dioxide is a direct result
1
of the greenhouse effect prevalent on the planet. This greenhouse effect is due to the planet’s close
proximity to the sun, a distance of roughly 0.72 AU. The surface temperature is an inhospitable 750 K and
the surface pressure about 90 atmospheres. Atmospheric density at the surface is one tenth that of water.
The surface atmospheric density of Earth is one thousandth that of water by comparison.
Another characteristic of the Venus environment is a layer of sulfuric acid found in the upper atmosphere.
This layer ranges from about 50 km to 60 km above the surface. Other cloud layers range from altitudes of
48 km to 68 km, with a layer of haze down to roughly 33 km. The atmosphere is clear beneath the haze
layer. Jet streams in the upper atmosphere travel with a speed of 85 m/s, circling the planet once every four
days. The motion of these jet streams is uniform resulting in little or no circulation. Winds on the surface
are much calmer, with speeds less than 3 m/s.
A successful Venus landing craft must be designed to withstand all elements of the Venus environment.
Previous missions such as Venera and Vega found that surviving for a significant length of time in such an
environment is a daunting task.
2
Chapter 2 - Mission Concepts
2.1 - Constrains
The mission constraints detailed by the AIAA competition Request for Proposal (RFP) include a minimum
sample return mass of 1.0 kg, the use of a US launch vehicle, and a budget limitation of 650 million dollars.
The mass of our design is directly limited by the US launch vehicle constraint coupled with the minimal
budgetary allowance. US launch vehicles are among the most expensive in the world, averaging between
100 and 200 million dollars per launch. The use of multiple launches is not practical due to this high cost.
The use of only one launch for this mission limits the mass of the entire design. Venus introduces its own
constraints through the harsh conditions on its surface. Temperatures exceeding 700 K and pressures up to
9 Earth atmospheres add a great deal of design complexity and structural mass to any system hoping to
survive on the Venusian surface.
2.2 - Propulsion Systems
2.2.1 - Orbiter
The orbiter’s propulsion system consists of a heliogyro solar sail design with counter spinning blade
segments. Each spinning segment has six blades attached to it. This counter spinning design removes the
angular momentum vector from the spacecraft to allow for steering and control through manipulation of the
sail blades. The heliogyro serves as the main propulsion system for the interplanetary, planetary capture,
and rendezvous portions of the mission. This concept requires no propellant mass to be taken on the trip to
and from Venus and allows for flexibility of launch dates and travel times.
2.2.2 - Venus Insertion Package
The Venus Insertion Package (VIP) uses a hydrazine and fluorine liquid propulsion system for de-orbiting
and controlling the Venus Lander during approach. The assumed Isp for the fuel to oxidizer mixture is about
425 seconds (Ref #19 p.692). Two thrusters and two spherical tanks containing the fuel and oxidizer for
each thruster are located on each axis. The VIP is capable of providing a 25 m/s ∆V along each axis for
attitude control and a 211 m/s ∆V along one axis for de-orbiting. The VIP separates from the Venus Lander
prior to ballute deployment.
2.2.3 - Venus Ascent Vehicle
The Venus Ascent Vehicle (VAV) is a two-stage solid propellant rocket. The propellant has an assumed Isp
of 290 seconds. The rocket is constructed from a graphite composite to maximize performance and
minimize mass. The first stage is launched from a cylindrical tank suspended from a balloon at an altitude
3
of around 61 km. The rocket thrusts vertically relative to the surface for five seconds and then begins a
gravity turn with an initial angle of 72 degrees. The first stage then burns for 75 seconds. The second stage
coasts along its trajectory for 530 seconds before beginning the final burn of 15 seconds. The final burn
places the sample capsule in an 800 km circular orbit around Venus.
2.2.4 - Earth Insertion Package
The Earth Insertion Package (EIP) uses the same basic system as the VIP. The insertion package is a
hydrazine and fluorine liquid propulsion system with two thrusters for each axis and two tanks for each
thruster. The EIP is designed to provide a total ∆V of 50 m/s for each axis and 1,500 m/s ∆V along one
axis for de-orbiting. The EEV is released from a 1,000,000-km orbit and requires more ∆V for attitude
control and de-orbiting. The EIP remains attached until it is jettisoned prior to Earth entry.
2.3 - Entry Systems
2.3.1 - Venus Entry
The Venus entry phase utilizes a ballute during atmospheric entry. The ballute deploys when the VIP is
released and the Venus Lander enters the appreciable atmosphere at an altitude of 180 km. The ballute
slows the lander to approximately 10 m/s at an altitude of 70 km. The upper aeroshell detaches from the
lower aeroshell, and the balloon is extended by this separation. The ballute and upper aeroshell remain
attached to the top of the balloon while the balloon inflates. Once the balloon is fully inflated it separates
from the aeroshell and the ballute and continues to descend. A few kilometers above the surface of Venus
the lower section of the aeroshell disconnects, allowing the legs of the lander to deploy.
2.3.2 - Earth Entry
The EEV is modeled after NASA and JPL’s Stardust Sample Return Capsule (Ref #23). The EIP thrusters
provide the ∆V to de-orbit and maintain the orientation of the EEV during Earth approach. The EEV enters
the atmosphere and free falls until drogue parachutes are deployed to slow it down so the main parachutes
can be deployed. A radio locator beacon is activated and the EEV continues to descend on the main
parachutes for a landing in the Pacific Ocean.
2.4 - Attitude Determination and Control Systems
The Attitude Determination and Control System (ADCS), like the power system, is divided into sections
corresponding to the three main segments of the spacecraft: the orbiter, the Venus Lander, and the EEV.
Each segment has its own ADCS because the segments operate separately from each other at various times
during the mission. The orbiter’s ADCS is the only one that is operational throughout the entire course of
the mission.
ADCS for the other segments become operational as necessary.
determination and control methods for each segment is provided in Table 1.
4
A summary of the
Table 1 - Attitude Determination and Control System Summary
SPACECRAFT SEGMENT:
Orbiter:
Venus Lander:
Earth Return Vehicle:
ADCS COMPONENTS:
Sun sensors
Star trackers
Hydrazine thrusters
Sun sensors
Star trackers
Hydrazine thrusters
Sun sensors
Horizon sensors
MANUFACTURER:
Ball Aerospace
Ball Aerospace
Ball Aerospace
Ball Aerospace
Ball Aerospace
Ithaco, Inc.
2.5 - Thermal
2.5.1 - Orbiter
The spacecraft’s thermal system tends the craft towards cold rather than hot. The system is designed this
way in order to prevent the components from overheating while in Venus’s orbit. A standard white paint
coating protects the antennas by increasing the reflection of solar radiation. While in transit to Venus and
back to Earth, cold sensitive components’ temperatures are regulated using Kapton heaters and multilayered insulation blankets. (Ref #2)
2.5.2 - Venus Lander
Thermal shielding is used for the rocket and instrument cylinders, and the sample container. The system is
based on a Multi-Layer Insulation (MLI) design. The outside layer is Ti-6AI-4V Titanium because of its
excellent strength to mass ratio, and its ability to tolerate the sulfuric acid found in the clouds of Venus.
The innermost layer is Type-304 stainless steel. A Fiberglass insulation and Xenon gas layer lies between
the inner and outer layers. The thickness of each layer is designed to withstand the atmospheric pressures
and temperatures of Venus.
2.6 - Mechanisms
The tables below describe the mechanisms used by the orbiter, the Earth return vehicle, and the Venus
lander. A brief description of each device, where each component is located, and the mass and the power
required by each is listed. More details are given in Sections 3.8, 4.7, and 4.8.
2.6.1 - Orbiter and Earth Entry Vehicle
Table 2 - Orbiter Mechanisms
Mechanism
Lightband separation
mechanism
Details
Detaches the Earth
entry vehicle (EEV)
from the main orbiter
Location
Interface between
orbiter bus and EEV;
Interface between
5
Mass
Power
1.363 kg
35 W
Thrusters
Sail fin motors
Magnetic bearings
EEV sample
collection cone
deployment
mechanism
Sample capture
"claws"
Sail fin base
separation mechanism
Communications dish
pointing mechanism
bus; Detaches the
Venus lander from the
EEV
Creates torque used to
deploy solar sail fins
Allows the fins to
rotate 180 degrees
Allows the fin bases
to rotate with no
friction
Spring loaded
telescoping
mechanism to deploy
cone from folded
position
Locks sample sphere
into EEV
Telescoping spring
separates fin bases
with enough room for
fins to rotate 180
degrees without
interference
Allows dish to rotate
and point in all
directions
Venus lander and
EEV
Two sets, one on each
fin base
At end of each fin
?
?
162 g
960
On beams connecting
the two solar sail fin
bases
On end of EEV next
to Venus lander
Inside EEV
0
310 kg
0
10 kg
0
49.2 kg
0
?
?
Between two solar
sail fin bases
Orbiter main bus
2.6.2 - Venus Lander
Table 3 - Lander Instruments
Instrument
Details
Graphite epoxy arm
with tungsten steel
lipped scoop attached
to the end
On side of
sample
cylinder
Corer
Designed and
manufactured by
Cybersonics Inc.
On bottom of
sample
cylinder
Variometer
Measures magnetic
Fields, or lack thereof
Within sample
cylinder
Wind Vane
Records consistency
of
wind direction
Top of lander
Panoramic
MicroImager
Acquires images of
Venusian surface and
of
sample collection
Within sample
cylinder
Mechanical
Arm
Location
6
Dimensions
0.5” innerdiameter
hollow tubes, two
6 ft segments
Mass
25 W
135
ma
x
g
15 cm long stem
×
2.67 cm
inner-diameter
45 cm tall
Power
1000 W
500 g
1W
250 g
2W
500 g
4W
Radar Altimeter
Measures altitude
Inside rocket
payload
container
4 kg
10 W
2.7 - Computer / Communications
2.7.1 - Computer
Many different versatile computer systems are required for this mission. The orbiter, Venus Lander, EEV,
and both insertion packages require computer systems to accomplish their necessary tasks. Each computer
system is capable of carrying out operations autonomously.
2.7.2 - Communications
The communication system for this spacecraft is designed around the fact that the majority of the mission is
completed autonomously. It is possible to update orbiter data to ensure that the attitude determination
sensors are as accurate as possible at all times. A steerable High Gain Antenna (HGA) is used to transfer
data at the maximum rate while allowing the spacecraft to remain on course during this transfer. Digital
cameras and omni-directional S-band antennas are used during the rendezvous phase of the mission. A
radio locator beacon is used on the EEV for sample location. NASA’s Deep Space Network (DSN) is used
to monitor the spacecraft during all phases of the mission.
2.8 - Rendezvous
The rendezvous phase is key to mission success. The orbiter actively tracks and intercepts the Venus
Sample Capsule (VSC) during this phase. Success of the rendezvous phase depends on insertion of the
VSC into close proximity of the orbiter. The VAV achieves its orbit and releases the VSC, which activates
the S-band radio beacon.
The orbiter, which is trailing several kilometers behind the VSC, locates the signal and closes in until the
VSC is within range of the digital cameras. The optical range of the cameras is approximately 100 m. The
cameras determine relative range, bearing, and range rate between the orbiter and the VSC. The orbiter
closes in on the VSC and catches it in the rendezvous cone. The VSC travels down the cone into the EEV
where three clamping mechanisms hold it in place. (Ref. #30)
2.9 - Power
The spacecraft power subsystem is divided into the following components: the orbiter, Venus insertion,
Venus Lander, Venus ascent, Earth insertion and Earth entry systems, each with its own power supply.
Each system has its own power requirements and must operate separately at various times during the
mission. Table 4 details the power budget for each segment of the spacecraft.
7
Table 4 - Power Requirements for Spacecraft Components
SPACECRAFT COMPONENT:
LANDER:
Computer
Drill
Arm
Deployment Mechanisms
Sensors
Sample Retrieval/Storage
VENUS THRUSTER PACKAGE:
ADCS Sensors
Computer
Thrusters
ORBITER:
Computer
Blade Motors
Thrusters
Antenna/Communications
Rendezvous Package
ADCS Sensors
ERV:
Locator Beacon
Parachute Deployment Mechanism
Computer
ERV THRUSTER PACKAGE:
ADCS Sensors
Computer
Thrusters
VENUS ROCKET:
Computer
ADCS Sensors
Ignition
POWER REQUIRED (Watts):
Total Power: 1043
15
1000
25
5
7
5
Total Power: 83
8
15
60
Total Power: 1120
15
960
5
60
65
15
Total Power: 25
5
5
15
Total Power: 83
8
15
60
Total Power: 25
15
8
2
Several variables affect the component selection and sizing of the power systems. Time is a major factor in
designing the power systems because the lifetime of the various components dictates whether or not
rechargeable batteries are necessary. Consideration is also given to the orbits maintained by the orbiter
segment because eclipse time about both Earth and Venus will affect the sizing of the rechargeable
batteries. Power requirements also vary with time depending on what components are operating and
whether or not they are constantly operating at peak power levels. Consideration is also given to the orbits
maintained by the orbiter segment, since eclipse time about both Earth and Venus will affect the sizing of
the rechargeable batteries. The individual power systems consist of various combinations of primary
batteries, secondary (rechargeable) batteries, and solar panels. Table 5 provides a summary of the power
sources used by each spacecraft segment.
8
Table 5 - Summary of Power Sources
SPACECRAFT SEGMENT:
Venus Lander:
Venus Thruster Package:
Venus Rocket:
Orbiter:
ERV Lander:
ERV Thruster Package:
POWER SUPPLY:
15 Li-Ion primary batteries
Li-Ion primary batteries
Li-Ion primary batteries
GaAs solar cells
NiCd rechargeable batteries
1 Li-Ion primary battery
Li-Ion primary batteries
9
MANUFACTURER:
Saft Battery Company
Saft Battery Company
Saft Battery Company
Spectrolab, Inc.
Sanyo Batteries
Saft Battery Company
Saft Battery Company
Chapter 3 - Main Orbiter Bus
3.1 - Configuration
3.1.1 - Heliogyro and Support Structure
The twelve blades of the heliogyro are made of a Kapton film approximately 2µm thick. The Kapton is
coated with a 0.5-µm thick layer of aluminum with a reflectivity of about 0.88 to 0.9. Each blade has a
length of 1,000 m and a width of 4 m. The blades taper to a width of 1 m at the root. This taper occurs
over a length of 4.45 m. Figure 1 shows a close up of a blade root.
Figure 1 – Blade Taper
The blades are set in a staggered configuration with two sets of six blades on each support ring. Figure 2
shows the staggered blade configuration. The overall length of the orbiter is 8.22 m in the stowed
configuration.
Figure 2 - Stowed Orbiter Configuration
10
The blades are connected to a hollow hexagonal support ring with a point-to-point diameter of 4.0 m. This
ring has a rectangular cross-section with a wall thickness of 0.635 cm, a depth of 0.3048 m and a width of
0.2 m. The panels of the ring structure are constructed of iso-grid aluminum. The center section of the
hexagonal ring structure is cut out to reduce the mass of the orbiter. Figure 3 shows the hexagonal support
ring.
Figure 3 - Hexagonal Ring Structure
The upper heliogyro support ring is rotated 3.5 degrees to ensure that the upper and lower blades do not
interfere with each other in their stowed configuration. The upper support ring is supported by aluminum
bars, which are used to stabilize the ring structure and carry the launch loads. The blades are connected to
the support rings by a blade arm. The blade arm is composed of a 1.0 m bar that is pinned to the center the
blade root. The blade arm is rotated and locked into the deployed position. Figure 4 illustrates the blade
arm deployment procedure.
Figure 4 - Blade Arm Deployment Procedure
The blade arms are connected to single-phase DC brushless motors within the heliogyro support ring.
These motors are used to rotate the blades to any angle desired. The main bus computer is used to calculate
this angle, which varies for each individual blade throughout the mission lifetime.
11
The blade deployment is separated into three stages: stowed, intermediate, and fully deployed. The
intermediate stage involves the extension of the upper heliogyro support structure. The blade arms then
rotated 90 degrees outward. The blades are unfurled and then rotated 90 degrees by the blade motors to
complete deployment. Permanent magnetic bearings are used to allow a frictionless rotation in both
support rings. Figures 5 and 6 show the intermediate and fully deployed stages respectively.
Figure 5 - Intermediate Orbiter Deployment
Figure 6 - Deployed Orbiter
See Appendix C for orbiter layout.
3.1.2 - Main Bus
The main bus is also a hexagonal structure with the same diameter, depth and thickness dimensions as the
heliogyro support rings. The center section is not removed to provide space for the computer, batteries, and
connection wires that branch out to the ADCS system and to the blade motors.
12
3.1.3 - Aeroshell
The aeroshell for the Venus Lander is composed of three aluminum conical sections and one aluminum face
section, each 0.635cm thick. Figure 7 is a rendering of the shell with dimensions included.
Figure 7 - Venus Lander Aeroshell
The upper section is the VIP, which also encases the ballute. The middle section houses the balloon and the
lander. The third section is the face of the aeroshell. The aeroshell is designed to protect the lander and
allow the hypersonic shock wave that forms around it to pass through the empty section of the deployed
ballute. The hypersonic shock profile was calculated using equation (3.1)
(3.1)
1
r
x
= 0.792 ⋅ C d 4 ⋅
d
d
from (Ref #1 p. 128). This equation uses the cross-sectional diameter of the face section (d = 4.15 m) and
the drag coefficient (Cd = 1.8) to calculate the distance from the centerline of the aeroshell to the hypersonic
shock wave (r) along the axis that is perpendicular to the blunt edge of the face section. Figure 8 shows the
profile of the hypersonic shock wave and Figure 9 shows the relative sizes of the aeroshell, hypersonic
shock wave and the ballute.
13
Hypersonic Shock Wave
12
Height to shock wave(m)
10
8
6
4
2
0
0
5
10
15
20
25
30
Distance from blunt edge(m)
Figure 8 – Hypersonic Shock Wave Profile
Figure 9 - Hypersonic Shock Wave Through Ballute
3.2 - Thermal
One side of the spacecraft is continually exposed to sunlight during the trip to Venus. All sides of the
spacecraft are exposed to extreme temperature gradients while in orbit around Venus. The solar intensity in
orbit at Venus is approximately twice the intensity encountered at Earth. The temperature of the outside of
14
the orbiter will reach as low as –200 degrees centigrade. The following thermal control system is modeled
after the Magellan spacecraft sent to Venus in 1989 (Ref #2).
The most sensitive components on the spacecraft are the electronics, batteries, and thruster propellant tanks.
Table 6 shows the operating temperature range for each.
Table 6 - Temperature Ranges for Sensitive Components (Ref #39)
Component
Battery
Charging
Discharging
Thruster propellant tank
Computer
Operating Temperature (°C)
Non-operating Temperature (°C)
0 – 45
-20 – 60
NA
-40 – 85
NA
NA
4 – 282
All electrical components and thruster propulsion tanks are wrapped in MLI blankets to protect them from
thermal extremes. The outside layer of the blanket is made of astroquartz, a material similar to glass-fiber
cloth that handles intense solar radiation extremely well. Chemical binders often used in astroquartz to
control flaking must be baked out to reduce the risk of discolorization leading to head buildup. The inner
layers of the blanket alternate between perforated, aluminized Mylar and B-4-A polyester netting. The
bottom layer is made of Kapton. The overall thickness of an average 8-layered blanket is 1.2 cm (the
netting is not counted in the number of layers). Figure 10 shows the layering of the thermal blanket used
on the Venus spacecraft.
Figure 10 - Multi-Layered Insulation Cross-Section
The antenna is coated with a white, inorganic, water-based paint developed at NASA’s Goddard Space
Flight Center. This paint reflects solar radiation and prevents discolorization. Electronic compartments in
15
the orbiter bus, Venus Lander, and EEV have louvers around them that open and close automatically to
regulate heat dissipation (Ref #2).
This thermal control system tends the craft and components toward cold temperatures. Kapton heaters are
therefore attached to protect cold-sensitive components.
Temperature sensors are mounted on each
component and software is written to ensure that the heaters are activated when a component becomes too
cold (Ref #2).
3.3 - Attitude Determination and Control System
3.3.1 - Attitude Determination
Control for the orbiter is performed solely by collective and cyclic pitch of the blades, so no additional
hardware such as control moment gyros or thrusters are required to provide pointing control. The attitude
determination sensors used by the orbiter are star trackers and sun sensors. The sun sensors are located on
the front face of the spacecraft, in the location most likely to maintain a position oriented towards the sun.
Two star trackers are located on the spacecraft bus, where spacecraft spin is not a factor. This combination
of sensors provides redundancy in the case of a sensor malfunction.
The star trackers are CT-602 High Accuracy Star Trackers provided by Ball Aerospace. These are small,
low mass devices capable of tracking up to five stars, with an accuracy of 3 arc seconds. Their Field-ofView (FOV) is approximately 7.8º ™ 7.8º, and they provide two-axis attitude determination as well as star
intensity data. They contain a radiation-hardened processor for environmental tolerance, and additional
memory for greater programmability. The sensor package also includes optics, a 512 ™ 512 pixel ChargeCoupled Device (CCD) detector, a thermoelectric cooler, command and data interface, and a spacecraft
power and mechanical interface.
The sun sensors are Ball Aerospace’s Precision Sun Tracking Sensors. These particular sun sensors have an
accuracy of 30 arc seconds with an 110º FOV, and are flexible for use on both spin stabilized and three-axis
stabilized spacecraft. The sun trackers, which are constructed of 6061 aluminum, are also radiation
hardened and use CCD based imaging. Each individual sun sensor is 0.165 m in diameter and 0.057 m tall
with a hexagonal cross section. The sun sensors, like the star trackers, provide two-axis determination.
3.3.2 - Control Systems
Spinning the heliogyro blades stabilizes them and removes the need for structure along the blades. The
spin rate is 0.04 radians/sec, or 2.3 degrees/sec. For a blade length of 1000 m, the corresponding angular
momentum is 39.79. The angular momentum vector points inward, normal to the face of the sail. The spin
of the sail provides the tension necessary to hold the blades flat and in the proper position. Rotating the
blades using the blade motors provides further attitude control. The blades are rotated in both collective
and cyclic manners.
16
Collective pitch constitutes applying a constant twist to each blade and is used to change the heliogyro spin
rate. The same torque must be applied to each blade for this maneuver, resulting in the same pitch angle for
each blade. This pitch mode is not time varying, because the angle applied is constant for all blades, and
does not change with rotation of the sail. This type of maneuver is also called a torque-control maneuver
(Ref #21 p. 88).
Cyclic pitch is time varying and may be modulated every rotation period. It is used to force the heliogyro
spin axis to precess by creating torques across the blade disk (McInnes, p.88). Pure cyclic pitch induces a
lateral force component in the plane of the blades that is used for planetary escape and capture spirals. Pure
cyclic pitch expressed as a function of time is given by equation (3.2).
(3.2)
θ (t ) = A sin(Ωt − ψ 0 )
A is the cyclic pitch amplitude component, V is the spin rate, and c0 is the phase angle. Pure cyclic pitch
contains no component of collective pitch. Cyclic and collective pitching can be combined for more
complicated maneuvers requiring both changes in spin rate and movement of the heliogyro spin axis. Such
maneuvers may be necessary for satisfying pointing requirements. This type of movement is used to
reorient the heliogyro and orbiter for capturing the VSC in Venus orbit. The equation of motion for this
mode is given by equation Error! Reference source not found..
(3.3)
θ (t ) = θ 0 + A ⋅ sin(Ωt − ψ 0 )
U0 is the collective pitch angle. Other, more complicated schemes can be generated to provide various
modes of control, depending on the mission requirements. The heliogyro can be fully controlled in all
flight modes and for all pointing requirements using various combinations of collective and cyclic control.
Determining the blade shape and coning angle from the spin rate and solar radiation effects is central to
controlling the heliogyro. A blade tensile stress analysis is performed on the heliogyro for this purpose.
For heliogyros with blade length R, chord C, and thickness h rotating with angular velocity Ω, the radial
and chordwise tensile stresses are determined by equations (3.4) and (3.5).
(3.4)
1
σ r ( r ) = ⋅ ρ ⋅ Ω2 ⋅ ( R 2 − r 2 )
2
(3.5)
1
C 
σ x ( x ) = ⋅ ρ ⋅ Ω 2 ⋅ [  − x 2 )
2
2
2
σr(r ) and σx(x) are the radial and chordwise tensile stresses, respectively (McInnes, p.88). The distance
outward along the sail blade is represented by r, and the distance in the chordwise direction along the sail
blade is represented by x. Results of these equations along the length of the blade are found in the
17
following graphs. The tensile stress decreases exponentially as one moves outward along the blade or
outward away from the midline at the root.
Figure 11 - Radial (Lengthwise) Stress Analysis
Figure 12 - Tensile Stress Along Blade Chord
The coning angle may be calculated after the tensile stresses for the blade are determined. The coning
angle is the blade curvature as a function of the distance from the root of the blade, and is expressed as
(3.6)
ϑ (r) =
2 ⋅ Pn
ρ ⋅ h ⋅ Ω2 ⋅ ( R + r)
18
where Pn is the solar radiation pressure for a given distance from the sun. Solar radiation pressure increases
as the orbiter gets closer to the sun. The following graph shows the variation in coning angle for a given
distance from the sun of 1.0 AU. The coning angle decreases with radial distance along the blade due to the
fact that the solar radiation pressure causes the blade to flatten.
Figure 13 - Coning Angle versus Position Along Length of Blade
It is possible to determine the blade shape as a function of radial distance once the coning angle variation
has been determined using the coning angle at the root, ϑ(0).
(3.7)
w(r ) = ϑ(0) ⋅ R ⋅ ln(1 +
r
)
R
Results of this equation for 0[r[R are graphed in the following figure for a distance from the sun of 1.0
AU. These results vary as the distance to the sun changes, because the solar radiation pressure varies.
19
1600
1400
Blade Shape
1200
1000
800
600
400
200
0
0
200
400
600
800
1000
1200
Position on Blade (m)
Figure 14 - Blade Shape versus Position Along Length of Blade
The blade twist, U, can be determined by solving the differential equation for blade twist as a function of r,
which is given by equation (3.8).
(3.8)
1
d 2θ
dθ
⋅ (R2 − r 2 ) ⋅ 2 − r ⋅
−θ = 0
2
dr
dr
This blade twist is independent of mechanical properties of the blade. A root torque Mo is required to twist
the blade through the desired angle and can be calculated using equation (3.9).
(3.9)
M0 =
1.208 ⋅ θ 0
⋅ I ⋅σ 0
R
U0 is the desired blade twist or pitch angle at the root, and s0 is the radial tensile stress at the blade root, and
I is the area moment of inertia, determined by equation (3.10).
(3.10)
I=
1 3
C h
12
It is obvious from the equation that the required torque increases linearly with increasing blade twist. The
torque is also small because the area moment of inertia is small due to the minimal thickness of the blades.
The results of this calculation are graphed in Figure 15 for pitch angles varying from 0 to 90 degrees.
20
Figure 15 - Required Torque versus Pitch Angle
3.4 - Power
The power system for the orbiter utilizes a solar panel to provide the required power. The cells used for the
solar panel are gallium arsenide cells with a germanium substrate, and are manufactured by Spectrolab Inc.
The cells are monolithic, two terminal, triple junction cells with a Beginning-of-Life (BOL) efficiency of
26% and an End-of-Life (EOL) efficiency of 21%. Each cell has an area of 30 cm , a thickness of 140 µm,
2
2
and a mass per unit area of 84 mg/cm (Ref #36). Assembly methods for the cells include soldering,
thermocompression, and ultrasonic wire bonding. The total area available for the solar panel, which is
2
located on the sun-facing surface of the orbiter, is 10.22 m . The power output for this area is 3,086 W.
This provides more than enough power for the orbiter.
The power generated by the solar panel is stored in secondary batteries for use during times of eclipse, such
as in Earth or Venus orbit. Nickel-Cadmium batteries provided by Sanyo are used as the secondary
batteries. The particular battery used is a KR-series CADNICA battery, the KR-10000M. The KR-series
Sanyo batteries are standard space-rated batteries known for their high performance and reliability. Data on
the Sanyo CADNICA battery chosen is provided in Table 7.
Table 7 - Orbiter Batteries
Sanyo CADNICA Battery (KR-10000M)
Nominal Voltage:
Capacity at 0.2C rate:
Diameter:
Height:
1.2 V
10000 mAh
43.1 mm
91.0 mm
21
Mass:
Charge Temperature Range:
Discharge Temperature Range:
Storage Temperature Range:
400 g
0oC – 45oC
-20oC – 60oC
-30oC – 50oC (-30oC – 35oC for long periods)
3.5 - Computer / Communication
3.5.1 - Computer
Calculating the motions of the individual blades in order to perform pointing, turning, and spin
maintenance is very complicated. The orbiter requires extensive computer calculations for ADCS. There
are two computers on the orbiter, one main computer and a backup computer. The computer selected for
the orbiter is the RHPPC Single Board Computer from Honeywell. This computer uses a radiation
hardened PowerPC 603e™ processor. This particular computer system is designed to operate for 15 years
in the severe thermal and radiation environments of space. Table 8 shows some of the features of the
computer system, and Figure 16 shows a picture of this computer system. Table 9 shows the radiation
hardness of the computer system. The software tools that can be used with this computer system are Wind
River Systems' Tornado™ environment, GNU C/C++ tools, and Wind River’s VxWorks™ realtime
o
operating system. The nominal temperature for this computer is 35 C, but it operates perfectly between -40
°C and 80 °C. The 210 MIPS provided by this computer is more than enough processor speed and power to
perform this mission. This system will cost about $400,000 for each computer (Ref #9).
Table 8 - RHPPC Feature Summary (Ref #9)
Processor
L2 cache
Memory
Backplane Bus
I/O
Debug/test port
Timers/counters
Form Factor
Mass
Power
Radiation hardness
RHPPC RISC (PowerPC 603e™ liscensed)
210 MIPS (Drhystone) @ 150MHz, 1.4 IPC
16Kbtye each Icache & Dcache
512KB, look aside, write through
4MByte SRAM, EDAC
4Mbyte EEPROM, super EDAC
64Kbyte SUROM (PROM)
cPCI, 32-bit, 33MHz, 3.3V
MIL-STD-1553B
2 Synchronous Serial full duplex ports, 12.5Mbps (RS422)
2 UART full duplex ports, 9.6K to 1M BAUD (RS422)
16 pins programmable as interrupt or discretes
JTAG (1149.1), COP, RHPPC debug
5, 32-bit general purpose, 4 with 8-bit prescale
50-bit mission timer with 6-bit prescale
32-bit watchdog timer, 2 stage
cPCI 6U x 220 (9.187” x 8.661”) with 2 PMC-like slots (74 x 149 mm)
2.2 pounds
12.5W (nom), 3.3Vdc ± 5%
Natural space
22
R
LM
111
R
R
R
512K x 8
SRAM
R
R
R
R
512K x 8
SRAM
R
512K x 8
SRAM
R
512K x 8
SRAM
R
1 553
D ual
XC V R
RH PPC
R
R
R
R
512K x 8
EE PR O M
512K x 8
EE PR O M
512K x 8
EE PR O M
51 2 K x 8
EE PR O M
512K x 8
EE PR O M
512K x 8
EE PR O M
51 2 K x 8
EE PR O M
512K x 8
EE PR O M
512K x 8
EE PR O M
512K x 8
EE PR O M
512K x 8
EE PR O M
R
OSC
O SC
512K x 8
SRAM
R
R
512K x 8
SRAM
R
RS422
Rcvr
26F 32
RS422
Drv
26LS 32
R
512K x 8
SRAM
R
512K x 8
SRAM
1553
XF M R
RS232
Rcvr
26F 32
PCI - PCI
B ridg e
1553
XF M R
R
R
LM
111
R
R
R
AC2 45
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
R
PPC-PEC
R
R
512K x 8
SRAM
R
R
R
R
R
512K x 8
SRAM
R
R
AC 245
R
AC 2 4 5
R
32K x 8
PRO M
R
AC 245
R
AC 245
AC 2 4 5
R
AC 2 4 5
32K x 8
PRO M
AC 2 4 5
512K x 8
SRAM
AC 2 4 5
512K x 8
SRAM
512K x 8
SRAM
512K x 8
SRAM
AC 2 4 5
512K x 8
SRAM
512K x 8
SRAM
512K x 8
SRAM
512K x 8
SRAM
PCI - PCI
B ri dg e
Figure 16 - RHPPC Mechanical Concept (Ref #9)
Table 9 - Radiation Hardness (Ref #9)
Total dose
Dose rate upset
Dose rate survive
Neutrons (1MeV DES)
SEU without L2
SEU with L2
Latchup
1E5 rad
1E8 rad/s
1E11 rad/s
1E13 n/cm2
4.4E-5 u/d
8.4E-5 u/d
none
3.5.2 - Communications
The steerable HGA utilizes an X-band signal with a frequency of about 80 GHz in order for the DSN
ground stations to pick up the data. The DSN is the monitoring agent on Earth for the duration of the
mission. DSN has the adequate coverage for the orbiter to be able to receive data at all times during the
mission. The HGA dish is 1.5 m in diameter and has a 2 m boom. The boom is connected to a 0.5 m arm
for stowing during launch. The HGA requires 60 W of power to operate and has a power output of 25 W
(Ref #11).
3.6 - Propulsion
The heliogyro propulsion system provides one main advantage over conventional chemical propulsion
systems. The mass of the propellant required to accomplish this mission with chemical propulsion is in
23
excess of 10,000 kg. The heliogyro propulsion system has a mass of roughly 300 kg by comparison.
Decreasing the overall launch mass reduces the size of the launch vehicle and ultimately the cost of the
mission.
3.6.1 - Solar Sailing Basics
A solar sail is a large, low mass reflective structure in space. Thrust is produced by photon pressure from
the Sun or other beamed energy sources. This concept gives a solar sail the ability to operate with an
unlimited supply of fuel (the Sun) within the inner solar system. When traveling in the outer solar system
the solar radiation pressure is significantly reduced (the drop in pressure falling proportionally with the
square of the distance to the Sun). Solar radiation pressure is the transfer of momentum from photons to
the sail. This momentum transfer occurs twice with the sail. The first transfer (Figure 17) occurs when the
photon strikes the sail, and the momentum of the photon is transferred to the sail/photon system.
Figure 17 - System before photon strike
The photon had momentum +p before making contact and the sail had 0 momentum and upon contact the
system has momentum +p. The photon is then reflected off the sail (the percentage of photons reflected
depends on the reflectivity of the sail material) transferring momentum to the sail (Figure 18). The photon
now has momentum –p and the sail has momentum +2p.
Figure 18 - System after photon strike
The optimal incident angle at which to hold the sail is 45 degrees with respect with the sun for a perfectly
reflecting surface. Surfaces do not reflect perfectly, which means that the angle of incidence is not equal to
24
the angle of reflection. The ideal pitch angle equals the cone angle of the reflected photon. Mathematical
modeling is done assuming an ideally behaving sail at 45 degrees.
3.6.2 - Equations of Motion
Patched conics are used to determine a preliminary trajectory for the solar sail. Circular planetary
motion and co-planar travel are assumed. The basic equation of motion in polar coordinates is
(3.11)
r
µ ⋅m r r
m ⋅ &r& = − 3 r + T
r
where µ is the gravitational constant of the influencing gravitational body, r is the
Figure 19 - Polar Coordinates Defined
position vector from the center of mass of the influencing body, T is the thrust acting on the sail, and m is
the mass of the spacecraft. Thrust for solar sail propulsion, T is given by equation (3.12).
(3.12)
r
 sin α ⋅ eˆr 
T = P ⋅ A

cos α ⋅ eˆθ 
where P is the solar radiation pressure, A is the surface area viewable by the sun and α is the angle of the
sail with respect to the Sun. Solar radiation pressure is given by equation (3.13).
(3.13)
P=
L(1 + ρ )
2 ⋅ R2 ⋅ c
25
where L is the solar luminosity, ρ is the reflectivity of the sail material, R is the distance to the Sun, and c is
the speed of light. Figure 19 illustrates the polar conventions used in these equations. Equation (3.14)
shows the second derivative of r in polar form.
(3.14)
(
)
(
)
r
r&& = &r& − r ⋅ θ& 2 ⋅ eˆr + r ⋅ θ&& + 2 ⋅ r& ⋅ θ& ⋅ eˆθ
Dividing equation (3.11) by the spacecraft mass and then combining equations (3.11), (3.12), and (3.13)
gives the acceleration of the solar sail as
(3.15)
&rr& = − µ rr + L(1 + ρ )A  sin α ⋅ eˆr 
r3
2 ⋅ R 2 ⋅ c cos α ⋅ eˆθ 
Equation (3.14) and (3.15) are equated and a system of two second order differential equations results when
the vectors are split into their components.
(3.16)
eˆr :
µ L(1 + ρ )A
&r& − r ⋅ θ& 2 = − 2 +
sin α
2 ⋅ R2 ⋅ c
r
(3.17)
eˆθ :
L(1 + ρ )A
r ⋅ θ&& + 2 ⋅ r& ⋅ θ& =
cos α
2 ⋅ R2 ⋅ c
Equation (3.16) and (3.17) are converted to a system of first order differential equations and solved
numerically using MatLab. These two equations make up the basis for all analysis done in computing the
trajectories for this mission. The front blades will periodically cast shadows on the rear blades because of
the counter spinning motion of the heliogyro. The shadows reduce the thrust output of the sail by reducing
the viewable area of the sail to the Sun. This was modeled as a simple sine curve that follows equation
(3.18).
(3.18)
 t ⋅ π ⋅ rpm 
T = .75 ⋅ Tmax + .25 ⋅ Tmax ⋅ sin

 180 
Tmax is the maximum thrust possible, which occurs when the sun strikes all the blades, t is the time in
seconds, and rpm is the revolutions per minute of the blades. Figure 20 shows the thrusting profile for the
first half of the mission (travel to Venus).
26
Figure 20 - Mean thrust for travel to Venus
3.6.3 - Interplanetary Travel
Assumptions are made in addition to those already mentioned above in order to simplify calculations to
obtain a reasonable estimate for the flight path of the spacecraft. The first assumption is to hold the solar
sail’s angle to the Sun, α, constant for the duration of the trip to and from Venus. The solar sail can vector
its thrust 90 degrees in either direction from the Sun. It is also assumed that the thrust produced by the
solar sail is the maximum possible at that distance from the Sun. The dynamics of this design allow
throttling from no thrust to full thrust. Both of these capabilities allow the solar sail to depart at anytime
and arrive at the desired destination without having to adhere to specific launch windows. Travel to Venus
is calculated holding the thrust angle at 135 degrees. Figure 21 shows the trajectory from Earth to Venus as
computed with these assumptions. The travel time is 452 days to reach the sphere of influence of the Venus
gravity field. This travel time is the minimum travel time from Earth to Venus using a sail area of 49,000
2
m and a payload of 2,912 kg.
27
Figure 21 - Travel Trajectory From Earth to Venus at Minimum Travel Time Conditions
Figure 22 - Travel Trajectory From Venus to Earth at Minimum Travel Time Conditions
To achieve this travel time a departure from Earth would have to be on March 28, 2005 at the earliest and
arrival at the Venus sphere of influence would occur on June 23, 2006. Launching on any other date
28
increases the travel time to intercept Venus. The return trip of the interplanetary travel phase (Figure 22) is
quicker due to the lower mass of the payload returning to Earth. The thrust angle for this portion of the trip
is held constant at 45 degrees in order to spiral outward. Again, considering the shortest possible trip time,
the first available launch opportunity comes on October 29, 2007. The travel time back to Earth would be
119 days and arrival at Earth would occur on February 26, 2008.
3.6.4 - Travel Around Venus
The travel within the sphere of influence of a planet was analyzed with a different set of assumptions from
that of the interplanetary travel. For this case the thrust angle is no longer held constant, rather the thrust
angle must vary with the position of the solar sail in the planet’s orbit. Energy is added or removed in order
to spiral out or in respectively. Thrust is vectored in the direction opposite the velocity vector to decrease
the energy and spiral towards the planet. When it is not possible to thrust against the velocity vector,
thrusting is stopped by positioning the blades of the sail parallel to the direction of the Sun. Figure 23
depicts this mode of thrusting involved in Phase I. When a suitable perigee is achieved in the capture orbit
a new approach to thrusting is taken in which thrusting only occurs within a specific distance from the
planet; in the case of the Venus capture orbit this radius is 100,000 km. This method of thrusting changes
perigee by 1000 km but reduces the apogee distance by over 100,000 km as can be seen in Figure 23 as
Phase II of the capture.
Figure 23 - Overview of Venus Capture
29
A closer inspection at perigee shows the small variance of the perigee point with each pass, shown in
Figure 24. Also shown is the entry path of the lander into the atmosphere.
Figure 24 - Venus Capture Close-up
For this case, Phase I lasts approximately 49 days and Phase II takes about 199 days. The orbiter releases
the lander at apogee just under four days after the completion of Phase II, and the lander takes 3 more days
to enter the atmosphere. Total time for this section of the mission is approximately 255 days.
The previous paragraph describes the latest design change in the mission. The original mission concept
called for the orbiter to spiral into an 800 km circular orbit to perform the rendezvous. Trajectory
computations calculating this orbit transfer produce unacceptable mission times of more than 3 years for
completion of the spiral transfer to the 800 km circular orbit. The need to change the mission to a more
appropriate time frame arose from this analysis. This new concept brings the orbiter into the highly
elliptical orbit shown in Figure 23, where perigee is at 800 km altitude and apogee is at 275,000 km. As
discussed in Section 4.10.2, the rocket would need to be resized to match such an orbit so that a rendezvous
can be achieved. With this mission concept the ∆V required to insert the lander into the Venus atmosphere
is reduced to 20 m/s at the apogee of the orbiter’s orbit (Figure 23).
The Venus escape portion of the mission begins after the rendezvous occurs. The escape is similar to the
capture Phase I in thrusting with the exception of that energy will now be added to the orbit by positioning
part of the thrust vector in the velocity direction. Figure 25 demonstrates an escape trajectory given
optimal conditions. This trajectory takes approximately 108 days to complete.
30
Figure 25 - Venus Escape Trajectory
3.6.5 - Future Analysis
There is room to improve the accuracy of these models. Matching the final conditions of interplanetary
travel with the initial conditions of planetary travel is one of the more important areas for improvement.
This is accomplished through vectored thrusting and throttling and would increase the estimated minimum
trip time on the order of months. Another area of improvement would be to 3-dimensionalize the
mathematics of the modeling.
The initial 2-D assumption made is reasonable, however, there is
approximately 3 degrees of inclination separating the Earth orbit from the Venus orbit around the Sun.
Referring to Figure 20, further study could be done to explain the flat portions of the thrust curve that occur
at the beginning of the transfer and at 300 days into the transfer.
Further analysis is also needed in the study of blade flutter that may result from the counter
spinning blades casting shadows on each other. This may or may not be a problem with the projected spin
rate of 0.38 revolutions/minute. Possible corrections to this problem include reshaping the blades to make
them thinner in width and longer in length or increasing the spin rate of the heliogyro. If the harmonic
response of this flutter is small enough and occurs away from the resonance frequency of the blades it may
also be an option to allow the small amount of flutter to occur. Other blade dynamics that are of concern
include the blade twist that can occur down the length of the blade and response lag to rotations applied at
the root.
31
3.7 - Mechanisms
3.7.1 - Lightband
Lightband is a separation device that allows two spacecraft to detach from one another. Walter Holemans of
the Planetary Systems Corporation invented the device’s simple design. A string held in tension holds the
two halves of the system together. When the tension is released by melting through the string, the two
halves separate and the spacecraft float apart. Lightband has heritage in the University Nanosat program
with the spacecraft trio of project ION-F involving the three universities: Virginia Tech, University of
Washington, and Utah State University. The system separated the three satellites successfully in 2003.
The Lightband mechanism allows the Venus lander to detach from the Earth return vehicle. In addition, the
separation of the Earth return vehicle from the main bus also utilizes Lightband. Lightband has a mass of
1.363 kg and requires a 30 W burst of power for activation. (Ref #13)
3.7.2 - Solar Sail Blade Thrusters
Thrusters along the outside of the orbiter fire to begin rotation of the solar sail bases. The rotation initiates
deployment of the solar sail blades. Once the correct rotational speed is reached, frictionless magnetic
bearings allow the rotation to continue without deceleration.
3.7.3 - Communications Dish Pointing Mechanism
A communications dish resides on the main orbiter bus where no rotation occurs. However, the dish itself
must be able to point in any direction within a full 360 degree circle. A motor will allow the dish to turn
and align itself for communication with Earth.
3.7.4 - Blade Rotation Motors
Blade rotation motors located within the blade support ring of the orbiter allow the solar sail blades to
rotate, enabling the sail to control the direction in which the spacecraft travels. Planetary gearhead motors,
model EC32, from Maxon Precision Motors Company, are used for the blade rotation. The motors provide
a torque of 2.25 N⋅m and have a mass of 162 grams. A total of twelve motors are needed to allow each
solar sail blade to rotate a full 180 degrees independently of other blades. Each motor requires 80 W to
operate. (Ref #20)
32
33
Chapter 4 - Venus Lander
4.1 - Configuration
The Venus Lander is mainly comprised of three large cylindrical tanks attached to a skeletal, titanium
platform. One of these cylinders houses the Venus Ascent Vehicle, and the other two contain the helium
necessary to inflate the entry ballute and the ascent balloon. Below the titanium frame, two smaller
cylindrical tanks contain the electronic, power and sample collection systems. The sample collection
cylinder housing connects to the rocket casing so that the sample can be inserted directly into the sample
capsule from below. Four telescopic legs support the lander and are attached to the main platform through
a pin and shock system. Figure 26 shows the lander in a fully deployed configuration.
Figure 26 - Deployed Venus Lander
The legs of the lander retract and are rotated up against the platform for the stowed configuration. Figure
27 shows the stowed configuration for the lander.
34
Figure 27 - Stowed Venus Lander
The shock absorbers slide along horizontal rails as the legs rotate, as shown in Figure 28, so they remain
extended throughout the descent until impact with the surface.
Figure 28 - Shock Absorber Deployed and Stowed Configurations
See Appendix B for lander configuration.
4.2 - Sizing Methodology
4.2.1 - Helium tanks
The layout of the lander is primarily based on the volume and mass of the top cylinders. The size of the
rocket housing came directly from the size of the Venus Ascent Vehicle. The materials used, and the mass
of the rocket housing is discussed in Section 4.3. Sizing of the two helium tank cylinders began with an
initial volume estimate based on the perfect gas law (equation 4.1) and the mass of helium needed for the
ballute and balloon of 70 kg. An internal temperature and pressure of 250 K and 19 MPa respectively were
3
used for this analysis. This application of equation 4.1 returned a volume of 2.1 m . This volume was cut
in half for lander symmetry. A hoop stress analysis (equation 4.2) was used to determine a wall thickness
for different combinations of vessel length and radius for this volume. Table 10 shows the results for
35
various geometric combinations for the half volume for four different materials. The highlighted row
represents the chosen geometry and material.
(4.1)
Pv = mRT
(4.2)
σz =
P⋅r
t
Table 10 - Helium Tank Geometry and Mass Combinations
Graphite
Vessel Radius (m) Vessel Length (m) Wall thickness (mm) Vessel Mass (kg)
0.1
33.29
1.38
45.09
0.2
8.09
2.76
45.73
0.3
3.31
4.15
47.47
0.4
1.56
5.53
50.86
0.5
0.67
6.91
56.45
0.6
0.13
8.29
64.79
Titanium
Vessel Radius (m) Vessel Length (m) Wall thickness (mm) Vessel Mass (kg)
0.1
33.29
2.07
197.76
0.2
8.09
4.13
200.61
0.3
3.31
6.20
208.32
0.4
1.56
8.27
223.34
0.5
0.67
10.34
248.10
0.6
0.13
12.40
285.04
Aluminum
Vessel Radius (m) Vessel Length (m) Wall thickness (mm) Vessel Mass (kg)
0.1
33.29
6.61
373.58
0.2
8.09
13.22
379.30
0.3
3.31
19.83
394.84
0.4
1.56
26.43
425.11
0.5
0.67
33.04
475.01
0.6
0.13
39.65
549.45
Steel
Vessel Radius (m) Vessel Length (m) Wall thickness (mm) Vessel Mass (kg)
0.1
33.29
1.38
229.83
0.2
8.09
2.76
233.10
0.3
3.31
4.15
241.98
0.4
1.56
5.53
259.26
0.5
0.67
6.91
287.74
0.6
0.13
8.29
330.24
36
The two helium tanks are constructed of graphite and include no heat shielding. Graphite was chosen
because of its high strength to density ratio. The ascent balloon is deployed during the descent to the
surface, so the integrity of the unshielded graphite tanks at the surface should not be an issue.
4.2.2 - Titanium Platform
The titanium platform is designed to withstand the maximum load occurring during the Venus entry phase
of the mission. The predicted 12 to 15-g deceleration creates a transverse load on all the bars in the
platform. These loads produce a maximum bending moment of about 6,200 N-m in the center cross-wise
bars Figure 29. The longest bars are used for the bending analysis because they experience the largest
bending load. We chose an initial outer diameter for the titanium bars of 5 cm. A required moment of
inertia for this load is determined using equation 4.3. The moment of inertia in this equation is then used to
solve for the inner diameter (equation 4.4). The thickness for these titanium bars is 1.1 cm. The mass for
the entire platform is 107 kg.
(4.3)
σz =
(4.4)
I=
M⋅y
I
π
4
4
⋅ ( ro − ri )
64
Figure 29 - Venus Lander Main Platform
4.2.3 - Landing Legs
The legs of the lander are telescoping, concentric cylinders. These legs are subjected to both bending and
buckling loads. In this case the bending load is the limiting factor. The shock absorbers are designed to
37
allow for 0.23 m of travel at an impact velocity of 4 m/s. This creates a bending load of 6,000 N·m in each
leg. This sizing includes a large factor of safety to account for possible difficulties with balloon inflation.
The upper/outer cylinder for each leg has the same dimensions as the bars used for the main platform. The
lower/inner cylinder for each leg is a solid 2.8 cm diameter bar. The entire leg is 1.6 m long when
deployed. These legs are pinned to the lander and lander feet, and they are connected to the shock
absorbers with ball and socket joints. The landing feet are 0.3 m diameter plates. Figure 30 shows a leg in
its deployed configuration.
Figure 30 - Venus Lander Leg Deployed Configuration
4.2.4 - Center of Mass
The center of mass for the lander is important both for aerodynamic stability during descent and to ensure
that the landing loads are distributed evenly across the platform. The upper cylinders are symmetrically
placed; the lower cylinders are not. The center of mass for the VAV is located 0.04 m from the center of the
platform. The sample collection cylinder is attached to the rocket casing directly below the payload
section, a distance of 0.75 m from the center of the platform. The rocket mass is 310 kg, and the mass of
the sample collection cylinder is 35.4 kg. The cylinder containing the battery pack has a mass of 53.1 kg.
Summing the moments about the center axis places the battery cylinder at a point 0.2 m from the center of
the platform on the same side as the rocket center of mass. Figure 31 shows the placement of these
cylinders with their respective distances from the center of the platform.
38
Figure 31 - Venus Lander Center of Mass Layout
4.3 - Thermal
The shielding is designed to prevent electronics meltdown premature rocket ignition. The rocket has a
safety temperature of 78 °C; at that point the fuel in the rocket ignites. The electronics must be maintained
at a temperature less than 40 °C. Figure 32 refers to pieces of the lander that use the thermal shielding for
protection.
Electronics
Container
Sample
Container
Figure 32 - Venus Lander Thermal Shields
The shielding is constructed of a variant of MLI comprised of a titanium (Ti-6AI-4V) outer shell of
thickness 38cm. Titanium was chosen because of its excellent strength-to-mass ratio and its ability to resist
39
sulfuric acid. The titanium is able to sustain the loads encountered during the Earth launch, Venus entry,
and Venus surface phases of the mission. The inner shell is comprised of Type-304 stainless steel with a
thickness of 0.76 mm. The purpose of the steel is to keep the insulation intact and against the wall of
titanium. Three sheets of micro-fiber felt composed of borosilicate glass fibers of thickness 1.3cm are used
for the thermal insulation between the inner and outer layers. Xenon gas located between the insulation
sheets is used with the borosilicate glass layer. Figure 33 shows the spacing and thickness of each piece of
the thermal shielding.
Figure 33 - Venus Thermal Shielding
The nominal heat transfer required for all the constraints to be sustained is 150 W. This is done using the
equation for heat transfer though a wall, equation (4.5).
(4.5)
Q=
k * (T2 − T1 ) * SA
L
Q is the heat transfer through the wall of the container, k is the thermal conductivity of the wall, T2 is the
outside temperature, T1 is the inside temperature, SA is the outer surface area of the cylinder, and L is the
thickness of the wall.
Figure 34 shows the insulation remains under 150 Watts for 3.33 hours at 460°C. This allows for ample
time to complete the sample collection phase of the mission. Figure 35 illustrates the thermal conductivity
at a variety of temperatures. (Ref. 12)
40
Heat Transfer (Watts) vs. Time (sec.)
160
140
Heat Transfer (W)
120
100
80
60
40
20
0
0
2000
4000
6000
8000
10000
12000
Time (sec.)
Figure 34 - Venus Shielding Heat Transfer versus Time
Thermal Conductivity vs. Temperature in Kelvin
0.045
Thermal Conductivity
0.04
0.035
0.03
0.025
0.02
0.015
0.01
0.005
0
250
300
350
400
450
500
550
600
650
700
750
800
Temperature (K)
Figure 35 - Thermal Conductivity versus Temperature
4.4 - Attitude Determination and Control Systems
The Venus Lander insertion segment and the EEV both use hydrazine fluoride thrusters for attitude control,
but the thrusters do not have the same characteristics for both systems. The Venus Lander is three-axis
controlled from separation with the orbiter to Venus entry. There are two thrusters along each axis for a
total of six thrusters. The total propellant mass is 20.371 kg. The fuel tanks have a radius of 0.099 m and a
thickness of 0.00043 m. The mass of each fuel tank is 0.149 kg, and the fuel mass is 4.074 kg. The
oxidizer tanks have a radius of 0.099 m and a thickness of 0.00044 m. The mass of each oxidizer tank is
0.151 kg, and the oxidizer mass is 6.111 kg. The thrusters provide a 25 m/s ∆V change per axis, or 12.5
m/s per thruster.
41
Attitude determination onboard the Venus Lander is accomplished through the use of a sun sensor and a
star sensor. Ball Aerospace provides both, and they are identical to the sensors used on the spacecraft
orbiter. Although they are housed in the lander insertion segment with the thrusters, they are located at a
distance away from the thrusters so that any propellant exhaust does not affect their performance.
4.5 - Power
The use of solar panels for power generation on Venus is impractical due to the lack of adequate sunlight on
the Venusian surface. Primary batteries are the best choice for a lander power supply; their number and
size are dictated by the lander power requirements. The Venus Lander requires three separate systems, one
for the lander, a second for the VIP, and a third for the rocket to transport the sample to the orbiter.
The lander must be operational for the duration of the surface mission in addition to the time required to
descend to the surface and then ascend to the appropriate altitude for rocket launch. Thus the lander
batteries must provide power for up to nine hours. Lithium ion batteries manufactured by Saft Battery
Company (Ref #32) provide an adequate power supply for the length of the surface mission, as well as for
the thruster package. The lander batteries are packaged as a cylindrical cell, in a sealed aluminum case, and
are affixed to the lander base. The thruster package batteries are packed the same way, but are contained in
the thruster housing section, and thus discarded before landing. Characteristics of Saft’s lithium ion
batteries are provided in Table 11.
Table 11 - Venus Lander Batteries (Ref #32)
SAFT LITHIUM ION BATTERIES:
Length:
Diameter:
Mass:
Average Voltage:
EOCV:
Power:
Specific Power:
Power Efficiency:
250 mm
54.2 mm
1,132 g
3.6V at C/2
4.1 V
132 Wh
117 Wh/kg
>94%
The lander power requirement for the surface mission is 1,043 W for 1.5 hours, for a total required capacity
of 1,564.5 W·hr. Additional power is required during the descent and ascent phases of the lander mission to
support the sensors and computer. Assuming a maximum time of 8 hours for the descent and ascent phases
of the Venus mission leads to a power requirement of 176 W·hr. These two power requirements together
yield a capacity of 1,740.5 W·hr. Using Saft’s lithium ion batteries to provide this power leads to a total
-4
3
battery mass of 14 kg, corresponding to 15 batteries each with a volume of 5.8x10 m , leading to a total
-3
3
battery pack volume of 8.12x10 m .
The Venus insertion package is operational from the time of release from the orbiter to the point at which it
is disconnected from the lander. The primary power requirement in this segment of the lander is the power
necessary to fire the thrusters. This is not a constant power requirement since the thrusters are not firing
42
continuously. The thruster package requires a power supply for five days, the length of time it must remain
operational. Assuming the thrusters are firing, the maximum power requirement is 83W. The power
requirement assuming no thrusters are firing is 23 W. Again, the lithium ion primary batteries will also be
used in this segment.
4.6 - Computer
4.6.1 - Venus Lander Computer
The Venus Descent Vehicle uses a Radiation Hardened Vector Processor (RHVP) by Honeywell. This
computer can handle 25 MIPS and has 2.7 MB of onboard static RAM. This system has a general-purpose
digital signal processor (DSP) onboard with the vector processor. Applications not suited for the fast vector
processor are handled by the DSP. The computer controls the separation of the Venus thruster package,
ballute deployment, balloon deployment, collecting and controlling scientific data while on the surface, and
transferring data to the sample capsule computer. Software supported includes Vector Builder, Vector Sim,
and COTS tools. This computer system costs about $100,000. (Ref #14)
4.6.2 - Sample Capsule Computer
The sample capsule computer will control the rockets flight, first and second stage separations, control the
omni directional s-ban antenna, and store all surface collected data. The sample capsule will be using the
RHVP computer along with a 48 MB extra solid-state memory card. The memory card cost about $200,000
and can store up to 48 MB of flash ram solid-state data. All of the data collected from the Venus Lander is
transferred from the lander’s computer to the sample capsule and stored in the extra memory card. This
entire system costs about $300,000. (Ref #14)
4.7 - Mechanisms
4.7.1 - Ultrasonic Drill/Corer
43
Figure 36 - Close up of the Ultrasonic Drill/Corer
An Ultrasonic Drill/Corer (USDC) acquires about 230 g of surface sample from Venus. The inner diameter
of the corer is 2.67 cm, and the drill stem length is 15 cm. The drill requires 1,000 W of power. The corer
works by using an ultrasonic horn driven at 20 to 23 kHz. A transformer converts the frequency to a drive
signal and a 60 to 1,000 Hz sonic wave. The USDC impacts the rock and creates fractures in the material
to achieve penetration.
Cybersonics Inc, located in Erie, Pennsylvania is working on the drill/corer in cooperation with NASA’s
JPL. The USDC can drill into hard rock samples including basalt, ice, and construction brick. This
instrument has the capability of drilling through Venus’s basalt surface to retrieve the required sample.
The corer connects to the bottom of the sample retrieval cylinder so that the drill stem tip hovers 0.35 cm
above the Venusian surface. An extend/retract device attached to the top of the corer allows the drill stem
to reach the surface. A ball and joint connection between the extend/retract device and the corer enables the
instrument to align itself perpendicular to the surface in case the lander does not land level.
The USDC bit is not sharpened, so there is no concern with the bit wearing out and losing performance
capabilities. The corer is ideal in that it is not subject to drill walk, does not apply large lateral forces on
its platform, and the drilling speed does not degrade with time. High temperatures are handled well by this
instrument as it has only two moving parts that are easily adapted.
(Ref #3)
4.7.2 - Mechanical Arm and Scoop
The mechanical arm will collect the rest of the 1.5 kg sample that must be returned. The arm requires a
power of up to 25 W. It is composed of two sections of graphite epoxy, each 0.4 m in length. Each section
is hollow, with a 2.54 cm inner diameter and a 3.0 cm outer diameter. The mass of the arm is 135 g, and
3
the scoop volume is 50 cm . The arm is composed of three sections: a shoulder joint joins the sample
retrieval cylinder and the first arm section, an elbow joint connects the two arm pieces, and a wrist joint
44
attaches the scoop to the arm. The rim of the scoop contains a top layer of tungsten steel to give the scoop
added toughness for trenching. Three narrow pieces of tungsten steel jut out from the back of the scoop.
These ripper tines (Ref #26) rake the surface to break up the rock and allow the scoop to collect the sample.
3
The arm digs up about 45 cm of Venusian dirt, tilts the scoop up, and allows the sample to fall down the
first section of the arm. A door connecting the sample container and the arm opens as the dirt begins to
descend down the arm. The pressure in the sample container is less than that on the surface, so when the
door opens a suction force helps collect the dirt falling down the arm. Meanwhile the mechanical arm
continues to dig and drop the samples through the hollow arm until a device in the sample container
informs the arm that the sample has been obtained. At this time the door to the sample container seals shut.
Through the procedure of using the suction force of the sample container, an atmospheric sample at surface
level is obtained in addition to the rock sample.
(Ref #26)
Lander
Sample Cylinder
Arm
Scoop
VENUSIAN SURFACE
4.7.3 - Sample Containers
The sample containers consist of a cylinder to hold the core sample, a surface sample sphere for the arm
sample, and two atmospheric sample spheres. The sample containers’ volumes were designed using the
density and characteristics of basalt. Extra volume was added to the arm sample sphere in order to also
provide room for atmospheric, surface level samples.
Two small containers for high altitude atmospheric samples are included. The containers are spherical in
shape and acquire their samples by suction. The computer signals the containers’ door mechanisms to open
at 70 km (above the cloud top) for one container and then at 40 km (in the cloud layer) for the other
container.
Atmospheric samples are obtained by using a pressure difference technique. The containers are pressurized
prior to Earth launch at a pressure lower than that of the sample to be collected. The container valve opens
at the desired altitude, allowing the atmospheric sample to be obtained. A timer is used to close the valve,
capturing the atmospheric sample in the container.
(Ref #)
45
4.8 - Scientific Instrumentation
4.8.1 - Variometer
A variometer is an instrument designed to detect magnetic fields. Venus is not believed to have a magnetic
field, but little data has been acquired in this area. The variometer is used to verify or disprove this belief.
The variometer requires 1 W of power and has a mass of 500 g. (Ref #8)
4.8.2 - Wind Vane
A wind vane is attached to the top of the lander. The wind vane deploys from the lander after the balloon is
deployed. This instrument determines whether the directions of the winds on Venus change or blow
constantly in one direction. The device is not designed to measure the magnitudes or direction of the
winds, but only the consistency of their direction. The wind vane requires 2 W of power and has a mass of
200 g.
4.8.3 - Panoramic Micro-Imager
The Panoramic Micro-Imager (PMI) acquires panoramic images of the Venusian surface and of the
mechanical arm and drill collecting surface samples. The PMI requires 4 W of power and has a mass of
500 kg. The imager is located in the sample cylinder under the lander. (Ref #8)
4.9 - Venus Entry and Descent
4.9.1 - Ballute Introduction
Entry designs were limited in the past by the mass cost of entry vehicles. The Venus Sample Return
Mission utilizes a new type of planetary entry device, known as a ballute, which allows for a reduction in
entry mass by the elimination of the massive heat shield.
A ballute is the physical union of a parachute and a balloon. The idea behind the ballute is to reduce the
heat flux incurred during entry until a heat shield is no longer necessary. A ballute achieves reduced
heating values by increasing the drag of the payload. A parachute cannot be used in the zero gravity
environment of space because of its inability to deploy. The ballute incorporates the rigidity of a balloon
with the drag characteristics of a parachute to increase the cross section and drag profile of the payload.
The ballute concept requires a drag producing and load carrying material, a storage and deployment
container, gas for inflation in zero gravity, and a tank to store the gas. The gas and tank requirements are
easy to fill because these systems are already provided for the ascent balloon.
46
4.9.2 - Shape
Ballutes can be either attached to the payload or connected by a tether. Attached ballutes, as seen in Figure
37, have the advantage of reducing the heating values on the payload to those of the ballute.
Figure 37 - Attached Aeroshell (Ref #18)
Detached ballutes can achieve higher drag profiles with smaller surface areas as evidenced by the lens
shape, which can be used to reach drag coefficient values as high as 2.0. The application of a disk with an
outer ring permits the further reduction of material usage while keeping the same essential shape and drag
profile. The outer ring is necessary to keep the peak heating values of the disk edge low. The ballute
chosen for the VSRM is in the shape of a torroid, as shown in Figure 38. The torroidal shape allows for the
shockwave from the payload to pass through the center of the ballute without affecting the flow around the
ballute. The material requirement for the ballute is minimized by using two tubes at each edge of the film,
as diagrammed in Figure 39. (Ref #22)
47
Figure 38 - Torroidal Ballute and Aeroshell
Figure 39 - Cross Section of Torroidal Ballute
4.9.3 - Materials
The materials required for this type of mission need to exhibit good thermal mechanical properties and
have high specific strength. A list of the possible materials for use in the ballute film application is
presented in Table 12 along with some of their properties. Fluoropolymers exhibit excellent thin film
48
qualities, but they are not considered due to their higher densities. Both Aramid and Kapton are produced
by
Table 12 - Ballute Film Materials (Ref #42)
Property
Density
Melting Temp
Glass Transition Temp
Tensile Strength
Tensile Elongation
Unit
g/cm3
°C
°C
kg/mm2
%
Kapton
1.420
none
350
18
70
Aramid
1.500
none
280
50
60
Polybenzoxazole
1.54
none
none
56-63
1-2
DuPont. Aramid is a long-chain synthetic polyamide, an organic thermoplastic (Ref #10), and Kapton is
a polyimide, a non-thermoplastic polymer. Aramid displays higher strength, but for this application the
thermal properties of Kapton are more desirable. Polybenzoxazole (PBO) is a liquid crystal polymer that is
developed by Foster-Miller. PBO demonstrates superior thermal and mechanical properties but is only
developed in smaller sizes making the fabrication of a large ballute difficult.
Kapton has a low density, intermediate thermal properties, and mature fabrication processes. Its lower
tensile strength is made insignificant by the addition of a load-carrying net around the ballute. Fibers for
the netting are presented in Table 13.
Table 13 - Ballute Fiber Materials (Ref #42)
Property
Density (g/cm3)
Tensile Strength (kg/mm2)
Elongation, Break (%)
PBO
1.56
577
3.0
Aramid
1.44
351-281
1.5-4.0
Spectra
0.97
306
3.5
Carbon
1.8-1.9
492-351
1.5-2.0
Spectra and Aramid both have poor high temperature characteristics, making PBO fibers the obvious choice
for use as the ballute netting. The tensile strength analysis of Kapton film and PBO fiber are shown in
Table 14.
Table 14 - Tensile Stress Analysis of Kapton and PBO (Ref #22)
Temperature °C
20
100
200
300
400
500
Kapton film (kg/mm2)
21.1
15.8
10.8
7.7
5.55
3.94
PBO fiber (kg/mm2)
577
473
363
254
208
200
The PBO fiber displays high strength even at high temperatures, and the Kapton film produces adequate
strength since it will not be the load carrying material. The structural integrity of the ballute throughout the
entire descent is not critical. The peak heating values occur in the matter of a couple minutes so any minor
holes ripped into the film have little effect on the success of the ballute as a whole.
49
4.9.4 - Sizing
Three factors influence the sizing of the ballute: the entry mass, entry speed, and material properties. The
entry mass, including the aeroshell, is 1,433 kg and the entry speed is 9.9 km/s. The maximum entry
temperature established by the ballute material is 500 °C, so a ballute radius of 23 m is required for a
material thickness of 14 µm. The 23 m radius represents the adequate radius for a disk shaped ballute,
2
creating a cross sectional area of 1,662 m . Manipulation of this data for a torroidal ballute is done by
determining the dimensions that create an equivalent surface area (Ref #22).
Optimization of the torroidal shape is accomplished by considering the material mass, mass of Helium gas
needed for inflation, and opening size needed for the shockwave. The Helium gas mass drives most of the
optimization because the large surface area of the ballute creates a large volume to be filled by the Helium.
An inflated pressure of 25 kPa is chosen because little pressure is needed to maintain the shape of the
ballute in the zero pressure, zero gravity environment of deployment. The results are presented in Figure
40 and Table 15. Final analysis shows that there is a 34 m opening for the shockwave to travel through.
Shortening the length of the connecting tubes that run between the lander and the ballute solves further
problems with the torroid swallowing the shockwave. The lengths of the fiber and tubing used are
approximated as suggested. (Ref #22).
Figure 40 - Ballute with Final Dimensions
50
Table 15 - Final Ballute Materials and Masses
Component
Film (Kapton, 20 g/m2)
Fiber (PBO fiber, 5 g/m)
Tubing (PBO film, 24 g/m)
Helium Gas (25 kPa)
Total
Size
1662 m2
15·Radius
20·Radius
145 m3
Mass (kg)
33.24
4
10
7
54.24
4.9.5 - Trajectory
Entry trajectory and profile analysis is accomplished using second order differential equations for the
position. The second time derivative of the position is formulated from equation (4.6).
v
v
v
F = m ⋅ a = m&x&
(4.6)
The position is calculated using a numerical integrator. The calculations take into account the drag and
gravity forces acting on the ballute. All other forces are considered negligible. The accelerations due to
drag and gravity are presented in equations (4.7) and (4.8) respectively.
r2
v C d 1 2 ρx& S
x&& =
m
(4.7)
v
µ v
x&& = 2 x
x
(4.8)
Cd is the coefficient of drag, which is only considered for the ballute because it generates most of the drag.
S is the cross section area of the ballute, and ρ is the density of the atmosphere. In the drag equation µ is
the gravitational parameter of Venus, 3.249 × 10 m /s .
14
3
2
An accurate density model is required for the Venus entry and landing phase of the mission, from the
beginning of the significant atmosphere to approximately 60 km. Equation (4.9) is obtained by curve
fitting atmospheric data (Ref #16).
(4.9)
ρ = e (-0.0000000912⋅x
4
+ 0.0000625⋅x 3 - 0.0133⋅x 2 + 0.8686⋅x - 17.35)
This density model is used to plot the planetary trajectory as shown in Figure 41. The velocity tangential to
the planet’s surface decreases rapidly, denoted by the sharp turn in trajectory. This plot leads to the
conclusion that the lander experiences large decelerations during descent. The deceleration loads can be
minimized using Figure 42. A minimum deceleration load of 7.7 g is possible at an entry angle of 5.65° for
an entry speed of 9.87 km/s. The entry sensitivity figure can be used to determine the entry angle corridor
51
that is necessary to sustain deceleration loads less than 10, which for the mission is approximately 0.2° and
gets smaller as the velocity increases.
Figure 41 - Venus Entry Trajectory
Entry trajectories are calculated using the conditions for the smallest deceleration load, and the resulting
deceleration and velocity values are plotted against the altitude in Figure 43 and Figure 44. Both of these
graphs show that the lander makes a slight skip during entry at an altitude of approximately a 115 km. The
velocity plot shows that the skip decreases the deceleration rate because the slope of the velocity curve
becomes larger at 115 km. These graphs also show that primary deceleration begins at approximately 120
-6
3
km altitude where the density is still on the order of 10 kg/m .
52
Figure 42 - Entry Sensitivity
Figure 43 - Entry Deceleration and Density vs. Altitude
53
Figure 44 - Velocity and Density verersus. Altitude
4.9.6 - Post-Entry Descent
The balloons multi-role mission includes descent and ascent. The most difficult situations for the balloon
during descent are extension (while connected to the ballute and lander) and then the subsequent inflation.
The balloon is extended by the ballute and upper aeroshell at nearly 70 km altitude while traveling at
approximately 10 m/s. The lander falls over 20 m as the balloon is extended and when the balloon is fully
extended the lander will undergo a nearly instantaneous 3 m/s change in velocity as the lander slows and
the ballute section speeds up again. This instantaneous load can be very dangerous to a thin balloon
material, so 5 rip-stitches are attached to the balloon to diminish the deceleration load. The rip-stitches are
located between the top of the aeroshell and the balloon and between the balloon and the platform. These
chords overlap the PBO fiber chords that also run between the balloon and platform, so that they can absorb
deceleration. Even if these chords melt during descent the PBO fibers are present to secure the attachment.
The descent trajectory uses the same equations and analysis techniques as the ascent trajectory. The initial
conditions are provided by the ballute entry trajectory analysis where the balloon is released at
54
approximately 65 km. The resulting analysis is presented in Figure 45 where balloon descent altitude is
plotted against time. Most importantly the lander touches down after nearly 3 hours of descent from the
upper atmosphere and the touchdown speed is only 4 m/s, as shown in Figure 46.
Figure 45 - Descent Altitude vs. Time
Figure 46 - Descent Velocity vs. Time
55
4.10 - Venus Ascent
4.10.1 - Venus Ascent Vehicle (Balloon)
The balloon is designed to serve two basic purposes; to slow the lander during descent and to raise the
rocket for eventual launch at an altitude of 60 km. The very limits of technology are pushed even farther in
designing a balloon to survive at such extremes as 150 mph winds, 460°C temperatures, 9 MPa pressures,
high impact velocities incurred while being dropped into the deployed position, and thick sulfuric acid
clouds. Balloon materials must exhibit:
•
•
•
•
•
•
•
•
low gas permeability
acceptable pinhole seaming
acceptable fabrication and folding
toughness in tear resistance
toughness in impact resistance
high specific strength
resistance to sulfuric acid
maintenance of mechanical properties at high temperatures
The Venus Sample Return Mission requires the lander to descend through the atmosphere to the surface,
acquire the sample, and then return to a higher altitude for rocket ignition. A balloon is very advantageous
to the Venus Sample Return Mission because it has uses in both the descent phase and the ascent phase.
4.10.1.a - Material Selection
Finding a balloon material to satisfy all of the mission requirements is nearly impossible. A list of possible
balloon materials is listed in Table 16, and comparing the data it becomes relatively obvious that there
really is no comparison. PBO, polybenzoxazole, is
Table 16 - Balloon material comparison (Ref #35)
Specific Strength (in)
Material
Film
Fiber
PBO
Teflon
HDPE
Kapton
Upilex R
Upilex S
Kevlar
Spectra 1000
Vectran HS
Nomex
68.8
1.5 – 4.0
4.5 – 5.9
9.6 – 20.6
34.3
64.8
495 – 516
16 – 30
519
392 – 458
602
345 – 432
93
Maximum Working
Temperature (°C)
500
260
80 – 120
250 – 320
270
290
180
147
110
310
a conjugated aromatic heterocyclic liquid crystalline polymer that can withstand the rigors of this
environment unlike no other known organic material. PBO’s rigid-rod molecular structure as shown in
Figure 47 creates a microscopic self-reinforcing structure that gives PBO the strength and stiffness of a
composite without the fiber and matrix interface problems. PBO has no melting temperature, no glass
56
transition temperature, and is highly resistant to corrosive chemicals. The rigid-rod structure makes PBO a
highly oriented fiber, which means that in its basic form, PBO will have little transverse strength. This is
commonly referred to as uniaxial orientation. This problem has been
Figure 47 - Chemical structure of PBO (Ref #42)
faced and solved with ingenuity by the people of Foster-Miller where a tri-modal die has been used to
produce biaxial PBO film. (Ref #35)
The mechanical properties of PBO also validate it for use in the Venus Sample Return Mission. Figure 48
displays the strength and modulus as they change with temperature
Figure 48 - Strength and Modulus vs. Temperature (Smith)
up to and beyond those temperatures that will be experienced during the mission.
The expected
temperature on the surface of Venus is 460°C, so Figure 2 shows that at nearly 50 and 100 degrees higher
than the maximum mission temperature PBO retains 36% and 28% of its strength respectively.
2
Remarkably PBO retains enough strength at 500°C (31 kg/mm ) to be stronger than Mylar, Kapton, and
PET at room temperature. (Ref #42)
57
Helium will be the buoyant gas selected for this mission due to its light weight; however, due to its small
size it is also important to ensure that the selected balloon material is not highly permeable to such a small
gas. The Helium permeability of the various candidate materials is shown in Figure 49, where once again
Permeability (mil/100 in
2
*atm *day)
PBO is shown to stand far above it’s
100000
10000
1000
100
10
1
0.1
PBO
Vectra
Polyester
Upilex
Peek
HDPE
Teflon
Material
Figure 49 - Helium Permeability of Several Possible Balloon Materials
competition as the best selection for balloon material.
Running sulfuric acid tests on PBO film shows that the acid has a profound effect on the mechanical
properties of PBO film. Samples soaked in sulfuric acid lose nearly 75% of their strength and become
plasticized. It was determined based on these results that a protective layer would need to be applied. (Ref
#31)
Possibilities for corrosive protection of the balloon include a fluoropolymer film and a noble metal coating.
Fluoropolymer films serve as a multilayer composite with the PBO. Fluoropolymers have very satisfactory
corrosive resistance, but they have poor heat resistance. The poor thermal-mechanical properties may be
excusable since the fluoropolymer would not be the load carrying film, but the other disadvantage is the
relatively high density of the fluoropolymers as shown in Table 17. Metal coatings can prove difficult to
sufficiently adhere to balloon materials but they also provide satisfactory corrosive protection for much less
mass. Based on the information provided in (Ref #31) from NASA, balloon manufacturers, and coating
experts, the best decision is to use a physical vapor disposition process to bond a protective layer of gold
onto the surface of PBO. For best adherence of the gold layer a tie coat metal layer would also need to be
deposited to the surface. Based on information from (Ref #31) the sufficient layering would be a tie coat
layer of Titanium from 50 to 100 Å thick and a gold layer from 1000 to 1200 Å thick. It is also noted that
the best adhesive results are obtained by a proprietary “heat and glow” treatment of the bonding surface.
58
The resulting mass, shown in Table 17, shows that the mass savings for metal coating would be drastic, and
makes the use of metal coatings an easy decision despite their tendency to crack upon folding.
Table 17 - Possible Corrosive Protection Materials
Material
Density (g/cm3)
Coating Mass (g/m2)
PFA film (12.5 microns)
2.13 – 2.16
26.6 – 27.0
PTFE film (12.5 microns)
2.13 – 2.20
26.6 – 27.5
FEP film (12.5 microns)
2.14 – 2.17
26.8 – 27.1
Gold layer (1200Å)
19.3
2.316
Titanium tie coat (100Å)
4.5
0.034
Total metal coatings
2.350
The balloon seaming is another important issue to be solved, since typical stitches, tapes, and adhesives
cannot be used to attach one PBO gore to another. The configuration of
Figure 50 - Balloon Seam (from 99-3858)
the balloon seam is presented in Figure 50. PBO is used to cover the seam since no tape is available that
has mechanical properties that are equivalent. The stitching material is graphite fiber. Currently the most
promising adhesive is a non-MDA condensation Avimid N based adhesive. Technological advancement
and research are necessary in the areas of adhesives, balloon seaming, and metal coatings to ensure success
of the mission.
4.10.1.b - Shape and Size
The type of balloon used for the Venus Sample Return Mission is a zero pressure balloon meaning that the
pressure inside and out of the balloon is equal. The vast range of pressures and densities from the surface
of Venus to the target altitude means that the volume of the balloon will change drastically throughout the
mission timeline. The sizing of the balloon is done in iterative steps considering the mass to be lifted,
desired altitude, balloon film thickness, and resulting size and mass of the balloon. The equations needed
for analysis of the balloon volume are the buoyancy force, equation (4.10), and the ideal gas law, equation
59
(4.11). Equation (4.12) is a manipulation of the buoyancy force equation, which gives the mass, mL, that
can be floated by balloon (for neutral buoyancy). V is the volume of the balloon, g is the acceleration due
to gravity, ρ is the density, R is the gas constant, and m is the mass of gas. Along with these equations it is
necessary to
(4.10)
F = ρgV
(4.11)
m
(4.12)
PV = mRT
L
= ρV
have equations for the temperature, pressure, and density for the range of altitudes that are to be utilized
during this phase of the mission, namely 0 to 66 km. The equations for temperature, pressure, and density
as a function of altitude in kilometers were derived using the data available in Venus and are presented in
equations (4.13), (4.14), and (4.15), respectively.
(4.13)
T = - 7.7992 ⋅ x + 731.88
(4.14)
P = exp(-0.0000101 ⋅ h 3 + 0.000088 ⋅ h 2 - 0.069 ⋅ x + 16.04)
(4.15)
ρ = exp(-0.000005 ⋅ h 2 - 0.0002 ⋅ h 2 - 0.0519 ⋅ h + 4.1772)
Initial size estimates for the balloon sizing are made using the mass of the rocket and casing alone, and a 51
micron thick film and are presented in Table 18.
Table 18 - Initial Balloon Sizing Analysis
Float Altitude
(km)
0
10
20
30
40
50
60
66
Balloon Float
Volume (m3)
6
11
20
41
95
264
910
2444
Required Mass of
Helium (kg)
38.2
37.7
37.4
37.5
37.6
37.7
37.8
37.9
Surface Area of
Sphere (m2)
17
24
36
57
101
199
454
878
Mass of Balloon
Film
1.3
1.9
2.8
4.5
8.0
15.7
35.9
69.3
It is apparent from this data that the balloon volumes needed to float at altitudes above 60 km begins to
grow exponentially. This process is done iteratively, while incorporating the values for the Helium mass
and film mass, realizing that the other masses will be minor in comparison to these issues.
A minimal required float altitude of 60km is established to finalize the calculations. A final balloon volume
is created by allotting for extra mass that is needed to complete the balloon, and accounting for extra gas
3
that creates lift above the required float lift. A fully inflated balloon volume of 1,375 m is established,
leaving the balloon shape as the next critical element. A sphere is the most efficient use of surface area for
60
a given volume, but the only concern for the sphere is the film stresses occurring during deployment and
partial inflation. Those reasons dictate the need of a cone-shaped lower section. The balloon shape used in
the VSRM has a spherical top and a truncated cone shape on the bottom, and the film thickness is 51 µm.
The equation used to determine the bottom shape is denoted by equation (4.16), where x is the radial length,
y is the height, R is the radius of the hemisphere, and m is a scaling factor. The final balloon shape is
2
shown in Figure 51 and has a surface area of 615 m .
 y 
x = R ⋅ cos 
 R ⋅ m 
(4.16)
The payload is attached to the balloon using four PBO film flaps (2 per balloon side) and chords made from
PBO fiber. Chords run from each film to locks on the four edges of the lander platform.
Figure 51 - Balloon with both payload attachments
These same chords also continue in towards the rocket casing where they are connected a second time so
that they serve in both phases. After the sample has been collected the locks at the edges of the platform
disengage and the chords become taught to the rocket casing. Upon completion of the landing phase the
rocket casing is released from the platform and the balloon rises to complete the second part of its mission.
4.10.1.c - Balloon Ascent
The final key to assuring proper balloon design is proving satisfactory ascent of the rocket. The ascent
analysis is performed using a numerical integrator to solve second order differential equations for the
position of the balloon. The accelerations on the balloon include gravity, drag, and buoyancy and their
equations are represented in equations (4.17), (4.18), and (4.19), respectively. The coefficient of drag for
the balloon is taken to be 0.9.
61
(4.17)
(4.18)
(4.19)
v
µ v
&x& = 2 x
x
r2
v C d 1 2 ρx& S
&x& =
m
v ρV µ v
&x& =
⋅
x
m x2
These equations are utilized along with the pressure, temperature, and density equations to model the
balloon altitude, velocity, volume, and times during ascent. The first important factor that ascent plays in
balloon design is the time to ascend. At 54 km the temperature is 39 °C, a temperature that does not
endanger the rocket. The time that it takes for the balloon to achieve that altitude is directly related to
Figure 52 - Lifting Gas Analysis
the amount of Helium gas that used in the balloon. Time for ascent decreases as the amount of Helium
increases, and conversely as the amount of Helium increases the maximum attainable altitude decreases.
This relationship is presented in Figure 52. The amount of Helium chosen is based on the storage tank
limitations (approximately 63 kg, maximum height 59.3 km), and once the float altitude is reached, 13 kg
of Helium is vented from the balloon to reach a top out altitude of 61 km. The final specifications are
shown in Table 19; the rip-stitches are discussed in the Venus descent phase.
Table 19 - Final Balloon Specifications
Material / Part
Balloon film, PBO film,
(51 µm thick)
Seams, Adhesives, and Stitches
62
Mass (kg)
48
5
Titanium tie coat (100Å)
0.02
Gold layer (1200Å)
1.4
Attachment Films
(PBO film)
Attachment chord
(PBO fiber, d=3.43mm)
4
0.200
Helium gas
63
5 Rip-stitches
5
Total
126.62
Using these values, the altitude is plotted against time for the ascent trajectory in Figure 53. It is important
to note that this data ignores the affects of high altitude winds. This analysis shows that the balloon ascent
takes just over 4.5 hours to reach maximum altitude and only 4.1 hours to reach 54 km.
Figure 53 - Ascent Altitude vs. Time
4.10.2 - Venus Ascent Vehicle (Rocket)
The Venus ascent vehicle is a two-stage rocket using solid propellant. The performance characteristics of
the propellant used to model the rocket are shown in Table 20. These characteristics are for a typical solid
propellant and were used for the preliminary design of the Venus ascent vehicle. In future studies, the use
of aluminized gelled propellants will be considered because they show an increase in performance for
rockets and have more complete, uniform propellant burns.
63
Table 20 - Propellant Performance Characteristics (Ref #17 p353)
Propellant Performance Characteristics
Chamber pressure (Pa)
5170000
Burning Time (s)
90
290
Isp (s)
3
Density (kg/m )
1800
Characteristic Velocity (m/s)
1527
The motor case design for the two stages is constructed of a graphite composite with an epoxy resin matrix
for structural stability. The properties of the graphite are listed in Table 21.
Other materials were
investigated during the design of the rocket, including Titanium, 2219 Aluminum, D6aC Steel, and 4130
Steel. The analysis performed on the Venusian rocket includes a 400 kg constraint on the total mass of the
system including payload since the rocket plus heat shielding needs to be lifted to a high altitude using a
balloon. Titanium was the only other material that had a positive payload of about 4 kg; the steel and
aluminum motor casings take up the entire dry mass payload. Graphite composite allows a payload size of
around 11.5 kg and so was decided upon as the material to use for the rocket even though graphite is the
most expensive material of the three.
Table 21 - Material Properties of Graphite (Ref #17 p310)
Graphite Properties
3
Density
(kg/m )
Tensile Strength (GPa)
Elasticity
(GPa)
1550
1.0
105
The modeling of the first and second stages was accomplished using the graphite and propellant properties
from Table 20 and Table 21. First, a single stage rocket was modeled to get an approximate size of the
rocket. The mass of the propellant, mp, was calculated using equation (4.20).
(4.20)
 Isp∆V⋅g

m p = m f ⋅  e 0 − 1




Eqatuion (4.20) contains a simple ∆V needed to achieve the desired orbit, plus a gravity and drag
correction. A 750 m/s (Ref #19 p722) drag correction was used along with a gravity correction of 3% (Ref
#19 p722) of the total ∆V. The final mass, mf, was specified at 20 kg to include unloaded rocket mass and
payload mass. This number kept the total mass of the rocket under the 400 kg constraint. Numbers are run
with these conditions for various materials as described above, and estimated propellant masses are
calculated. Simple geometry equations are used to calculated the motor case size and estimate the unloaded
rocket mass. The thickness of the motor casing is estimated using a burst pressure and a factor of safety of
1.25. Once the initial rocket design is completed, a more detailed launch profile is done to determine the
number of stages and the size of each stage necessary to achieve the desired orbit. The detailed launch
profile is explained later in this section. Two burns are needed to achieve the desired orbit. The initial
64
estimations of gravity and drag prove to be good since there is just the right amount of propellant to get into
the desired orbit. The time for each burn determines how the propellant is divided between the two stages.
These propellant masses are then input back into the rocket analysis to design the motor case and size for
each stage. Table 22 shows the details of the first stage of the Venus Ascent Vehicle. To maximize the
thrust, the nozzle is designed to expand to the average pressure that the first stage experiences. The throat
area of the nozzle is determined using equation (4.21) (Ref #17 p310). Exit areas of the nozzles are
calculated using equation (4.22) (Ref #38 p55).
Table 22 - Venus Ascent Vehicle Stage One Configuration
Mass Properties
(kg)
242
6
0.0002
0.0005
248
Propellant
Motor Case
Nozzle
Igniter
Total
Motor Case Properties
Length
Radius
Thickness
(m)
1.3
0.18
0.0016
Rocket Nozzle Properties
2
Throat Area (m )
0.00095
2
Exit Area (m )
0.0066
Throat Diameter (m) 0.035
At = CV
(4.21)
(4.22)
Ae =
mp
tb ⋅ PC
At
1
 γ + 1  γ −1  Pe

 ⋅
 2 
 PC
1
γ
γ + 1  Pe 
 ⋅
⋅ 1 − 
γ − 1  Pc 

γ −1
γ
Table 23 shows the details of the second stage of the Venus Ascent Vehicle. The nozzle for the second
stage is modeled to bring the exit pressure close to zero since the second stage fires outside the Venusian
atmosphere. Figure 54 is an AutoCAD rendering of the Venus Ascent Vehicle. The overall length of the
entire rocket is 2.15 m from the bottom of the first stage nozzle to the top of the blunt nose cone. Figure 55
shows the heights of the motor casing and nozzles with the sample capsule in the payload area. This rocket
design allows for an 11.5 kg sample capsule payload. The use of a better propellant increases the payload
mass; this is an area needing further study for optimization. The rocket’s total mass is about 310 kg
including the mass of the loaded sample capsule.
65
Table 23 - Venus Ascent Vehicle Stage Two Configuration
Mass Properties
Propellant
(kg)
49
2.5
-05
8.00×10
0.00019
51
Motor Case
Nozzle
Igniter
Total
Motor Case Properties
Length
Radius
Thickness
(m)
0.26
0.18
0.0016
Rocket Nozzle Properties
2
Throat Area (m )
2
Exit Area (m )
Throat Diameter (m)
0.00095
0.0115
0.035
Figure 54 - Venus Ascent Vehicle Concept
66
Figure 55 - Venus Ascent Vehicle Dimensions
The rocket’s flight profile is calculated using the equations of motion that take into account a drag profile
and the gravity turn as seen in equation (4.23).
r
r
r r
mr&& = Thrust + Drag + Gravity
(4.23)
Thrust was modeled using equation (4.24). Drag was modeled using equation (4.25) and the gravity was
modeled using equation (4.26). The symbols used are defined as follows: γ is the flight path angle, Ve is the
exit velocity, ρ is the density, A is the cross sectional area, V is the current velocity, CD is the drag
coefficient, µ is the gravitational constant, r is the current radius from Venus, and er and eθ are the
coordinate system of the rocket.
(4.24)
r
Thrust = Ve ⋅ m& ⋅ (sin(γ ) ⋅ eˆr + cos(γ ) ⋅ eˆθ )
(4.25)
r
1
Drag = ⋅ ρ ⋅ V 2 ⋅ C D ⋅ A ⋅ (sin(γ ) ⋅ eˆr + cos(γ ) ⋅ eˆθ )
2
(4.26)
r
µ ⋅ mrocket
Gravity =
⋅ eˆr
r2
These equations are turned into four first order ordinary differential equations and solved using MatLab.
Initial conditions are modified until the desired orbit is attained. The following data are the results from an
800 km circular orbit optimization. The first stage provides 75 seconds of constant thrust getting the rocket
to 147 km off the surface of Venus. Next, the first stage motor case and nozzle are ejected leaving the
67
payload and second stage. The rocket remains in an elliptical orbit for 534 seconds. After 534 seconds the
rocket becomes tangent to the desired 800 km orbit. The second stage then fires for 15 seconds bringing
the speed of the rocket to match that of the 800 km desired orbit. The second stage motor case and nozzle
are then ejected from the payload sample capsule. The payload sample capsule remains in orbit until the
orbiter rendezvous and collects the capsule. Figure 56 shows the over all launch profile from the surface of
Venus to the 800 km orbit. Figure 57 is a close up of the surface to orbit launch profile. The profile of the
rocket’s altitude with respect to time can be seen in Figure 58.
The orbiter is unable to achieve an 800 km circular orbit within three years. In the next design loop the
rocket needs to be resized for an elliptical orbit with a perigee of 800 km with a velocity of 9.53 km/s and
an apogee of 273,000 km. Initial calculations determine that the rocket only needs 2.9 seconds of extra
thrust during the second stage to achieve this orbit.
Figure 56 - Venus Ascent Vehicle Flight Path Profile
68
Figure 57 - Venus Ascent Vehicle Launch Profile
69
Figure 58 - Venus Ascent Vehicle Altitude versus Time Plot
70
Chapter 5 - Earth Entry Vehicle
5.1 - Configuration
The EEV is a 60-degree blunt body capsule, which is modeled after NASA and JPL’s Stardust Sample
Return Capsule (Ref #23). The EEV is composed of aluminum with a maximum cross sectional diameter
of 1.5 m and a height of 0.9 m. Two drogue parachutes and three reinforced ring-slot descent parachutes
are contained inside the EEV (Ref #24). Barometric switches are used to deploy the parachutes at the
appropriate altitude (Ref #28). Also contained inside the EEV are the onboard computer, batteries, and a
radio locator beacon. The radio locator beacon is activated upon parachute deployment. The radio locator
uses approximately 5 W of power and has a range of 3 km.
The Venus Sample Capsule remains in the captured position throughout the entry and landing procedure.
No floatation system is used for the EEV. The EEV displaces 955 kg of seawater when fully submerged,
and the mass of the EEV is 242 kg, therefore the EEV is buoyant in water.
5.1.1 - Sample Collector
It is the job of the orbiter to capture the VSC after the VAV launches it into orbit. A rendezvous cone,
which is modeled after the European Space Agency (ESA) rendezvous cone, is extended when the Venus
Lander detaches from the orbiter (Ref #8). The rendezvous cone is designed so that once the VSC enters
the cone it cannot bounce out. The narrow end of the capture cone leads to a narrowing cylinder that runs
through the Earth Entry Vehicle. The VSC travels through the EEV cylinder to the narrow end where it
hard-docks with the EEV.
Figure 59 - Orbiter, EEV, with Extended Cone
71
The VSC travels through the EEV cylinder and comes to rest in the Capsule Containment Compartment.
Three capture claws latch around the VSC to hold it in place until it can be retrieved after Earth entry (Ref
#8). After the sample capture maneuver is complete, the rendezvous cone is jettisoned to reduce the
spacecraft mass for the return trip to Earth.
5.2 - Thermal
The heat shielding for the EEV consists of two layers of Composite Flexible Blanket Insulation (CFBI).
CFBI is designed to have high emissivity, resistance to a high heat flux, and insulation capability between
1450 and 1650 degrees C. Two types of insulation, CFBI-1 and CFBI-2, have been tested at pressures of
1.0, 0.1 and 0.01 atm and temperatures of 23, 200, and 400 degrees C (Ref #41).
The apparent thermal
conductivity of CFBI-1 and CFBI-2 is shown in Table 24.
Table 24 - Apparent Thermal Conductivity (Ref #41)
Temperature
(°C)
at 1.0 ATM
23
200
400
CFBI-1
CFBI-2
Thermal Conductivity
(W/m⋅K)
0.036
0.035
0.051
0.049
0.059
0.084
CFBI-1 is made of silicon carbide and alumina covered with aluminized Kapton on one side, whereas
CFBI-2 is covered with aluminized Kapton on both sides. Both CFBI-1 and CFBI-2 are 0.026 m thick and
3
3
have average densities of 133 kg/m and 149 kg/m respectively (Ref #41). CFBI-1 is chosen over CFBI-2
for use on the Earth Entry Vehicle because of its lower average density and lower thermal conductivity at
higher temperatures.
5.3 - Attitude Determination and Control Systems
The EEV requires a separate ADCS from the orbiter, since it operates independently during Earth re-entry
and landing. The EEV is equipped with hydrazine-fluoride thrusters for three-axis attitude control. There
are two thrusters directed along each axis, for a total of six thrusters. The tanks required for the fuel have a
radius of 0.059 m, thickness of 0.00026 m, and mass of 0.0314 kg. The fuel mass is 0.858 kg. The
oxidizer tanks are the same size, and the oxidizer mass is 1.287 kg. The thrusters are fired when needed to
maintain the proper orientation in orbit and for re-entry. Each thruster provides a DV of 25 m/s.
The EEV has two types of attitude determination sensors: sun sensors and horizon sensors. The sun sensor
used is identical to those on the spacecraft orbiter. The horizon sensor is the Horizon Crossing Indicator
(HCI) provided by Ithaco. The HCI provides 0.1-degree accuracy with a 1.0 degree x 1.0 degree FOV, with
a low mass of 0.65 kg. It requires less than 0.7 W of power with a peak current of 3 A. Both sensors are
mounted on the EIP, but at a distance away from the thrusters such that any propellant exhaust will not
affect their performance.
72
5.4 - Power
The EEV has two separate power systems, one for the return vehicle, and one for the attached EIP. Lithium
Ion batteries from Saft are utilized, with the battery packages resized to fit the requirements of the EEV
lander and EIP. The two power systems are separate since the EIP is discarded once atmospheric reentry is
underway, and the EEV still requires power for the beacon and parachute deployment mechanism. The
power capacity required for the EEV primary battery is approximately 25 W·hr, assuming a 15 minute time
period for atmospheric entry and additional operational time the EEV to be detected and found. One
lithium ion primary battery is sufficient for this final phase of the mission.
5.5 - Computer
The earth lander computer will control the thruster package separation, the parachute deployment schemes,
and control the radio locator beacon that is located on the lander.
The Dual Single Board Computer (DSBC) was selected for this mission. This computer system can
perform 1.15 MIPS at 6 MHz and only requires 5.5 W of power at its maximum peak and 4 W of power
during normal operation. This computer system is radiation hardened, costs nearly $50,000, and can be
seen in Error! Reference source not found.. (Ref #14)
Figure 60 - DSBC Computer
5.6 - Propulsion
The EIP consists of two main thrusters, which are used to provide the necessary ∆V to insert the EEV into a
direct Earth entry trajectory. The EEV attitude control system consists of four secondary thrusters, which
are used to re-orient the EEV during the Earth entry insertion maneuver and to maintain the proper attitude
for entry. Each thruster on the EIP utilizes a fluorine (F2) and hydrazine (N2H4) bi-propellant system.
73
Fluorine and hydrazine bi-propellant systems are used because of their relatively high vacuum specific
impulse (Isp) of 425 seconds, high thrust range of 5 N to 5×10 N. These systems also have low average
6
3
3
oxidizer and fuel bulk densities of 1.5 g/cm and 1.0 g/cm for fluorine and hydrazine respectively (Ref
#SMAD p692).
Table 25 - Performance Characteristics of Propulsion Systems (Ref #19 p692)
Type
Cold gas
Liquid:
Monopropellant
Bipropellant
*
Propellant
Energy
N2, NH3,
Freon,
Helium
High
pressure
H2O2, N2H4
Exothermic
decomposition
Chemical
Chemical
Chemical
Chemical
O2 and RP-1
F2 and N2H4
CIF5 and N2H4
N2O4/N2H4
Dual Mode
Gas densities at STP
Vacuum
Isp
(sec)
50-75
Thrust
Range
(N)
0.05-200
Avg. Bulk
Density
(g/cm3)
0.28*, 0.60,
0.96*
150-225
0.05-0.5
1.44, 1.0
350
425
350
330
5-5×106
5-5×106
5-5×106
3-200
1.14 and 0.80
1.5 and 1.0
1.9 and 1.0
1.9 and 1.0
5.7 - Earth Entry and Descent
The orbiter releases the EEV at the edge of Earth’s sphere of influence The EEV is facing in the orbital
velocity direction, and the heat shield is facing in the negative orbital velocity direction when the EEV
detaches from the orbiter.
Figure 61 - Orbiter and Earth Entry Vehicle Separation.
74
The EIP thrusters fire to provide the necessary ∆V to insert the EEV into a direct Earth entry trajectory.
Next, the EIP thrusters fire to rotate the EEV 180-degrees about its non-symmetry axis so that the heat
shield is facing in the orbital velocity direction to prepare for atmospheric entry. The EIP is separated prior
to entry. The EEV enters the Earth’s atmosphere at approximately 11 km/s with an entry angle of
approximately 10 degree, and then free falls through the atmosphere. Two drogue parachutes are released to
slow down the EEV from approximately 140 m/s to 80 m/s at approximately 7 km altitude.
At
approximately 3 km altitude three main descent parachutes are deployed and an omni-directional radio
locator beacon is activated. The main parachutes decelerate the EEV to a final landing velocity of
approximately 9 m/s (Ref #28). The VSC is recovered and the sample is taken to the appropriate facility
for analysis.
Figure 62 - Earth Entry Vehicle with Descent Parachutes
5.8 - Sample Analysis
Electron-microprobe, X-ray diffraction (XRD), transmission electron microscope (TEM) and scanning
electron microscopes (SEM) are techniques commonly used to obtain the mineral composition. All of these
techniques can be done in house at a university such as Virginia Tech. X-ray fluorescence (XRF) analysis
will be done to determine the bulk composition of the Venus sample. Virginia Tech does not have the
equipment to perform this bulk analysis, so it will have to be done at a national laboratory
The cost of all analysis done on the Venus sample depends on scientist salaries, equipment cost, sample
preparation time, and sample analysis time. Sample preparation time may take several hours, but sample
analysis time varies widely. Dr. Benedix of the Virginia Tech Geology Department estimates that the entire
sample analysis will cost approximately $50,000 (Ref #5).
75
Chapter 6 - Cost Analysis
Determining the cost of a mission of this scope is inherently difficult. Many manufacturers are hesitant to
give out actual cost numbers for materials and fabrication for design projects such as this. The prices and
fabrication costs for some components are presented in Table _. Many of the structural components and off
the shelf electronics are presented here. Fabrication and material costs for experimental components such
as the insertion ballute, ascent balloon, and the aluminum coating for the sail blades are not known. The
total known cost for the mission of approximately 150 million dollars is well below the 650 million-dollar
budget. This cost estimate does not include any ground support costs or developmental costs for the
experimental components.
76
Table 26 - Component Costs
Component
Computer
(orbiter)
lander
sample case
Earth lander
Memory
(sample)
Utrasonic Drill
Mechanical
Arm
Sample
Container
Sun sensors and
star trackers
Delta IV
Helio-gyro
blade Kapton
layer
Company
Purchasing
From
Description
Cost/Unit
# of units
Total
603-E
RHCP 32
vector chip
RHCP 32
vector chip
RHCP 32
vector chip
Honeywell Co.
$400,000.00
2
$800,000.00
Honeywell Co.
$100,000.00
2
$200,000.00
Honeywell Co.
$100,000.00
1
$100,000.00
Honeywell Co.
$100,000.00
1
$100,000.00
45 MB
Honeywell Co.
Cybersonics
Inc.
$200,000.00
1
$200,000.00
$8,000.00
1
$8,000.00
hollow graphite
epoxy
in house
$100,000.00
$10,000.00
M+
Ball Aerospace
Boeing Co.
less than
$100,000.00
1
$100,000.00
$145,000,000.00
1
$145,000,000.00
$26,333.33
12
$316,000.00
DuPont
Table 27 - Fabrication Costs
Component
Venus lander
platform
Lander legs
Heliogyro rings
Orbiter bus
Aeroshell
Description
Company
Purchasing From
Only structure
10 days at large
facility
Cost/Unit
# of units
Total
$1,200.00
$1,000.00
$4,000.00
$4,800.00
1
4
2
1
$1,200.00
$4,000.00
$8,000.00
$4,800.00
$32,000.00
1
$32,000.00
The cost of all analysis done on the Venus sample depends on scientist salaries, equipment cost, sample
preparation time, and sample analysis time. Sample preparation time may take several hours, but sample
analysis time varies widely. Dr. Benedix of the Virginia Tech Geology Department estimates that the entire
sample analysis will cost approximately $50,000 (Ref #5).
77
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Appendix A – Mission Timeline
82
Appendix B – Venus Lander Schematic
83
Appendix C – Orbiter Schematic
84