Payload Concept Proposal
Mercury Lander Mission
Spring 2012
Cold Springs High School
Team 1
The F.O.R.C.E.
Payload Concept Proposal
Mercury Lander Mission
Spring 2012
1.0 Introduction
As Mercury orbits the Sun, it is continually bombarded by solar wind. The solar wind is a supersonic
outflow of plasma into space. From this extreme form of energy particles come varying velocities,
temperatures, densities, and magnetic field properties. As these varying properties collide, co-rotating
interaction regions (CIRs) are formed. CIRs are essentially interplanetary heat tornadoes. The F.O.R.C.E.
(Flying On Resources from Continuous solar wind Energy) has developed a payload design called Yoda.
Yoda is designed to determine the force, acceleration, and pressure of the solar winds.
2.0 Science Objective and Instrumentation
The general concept of this payload is to be launched from the spacecraft on its approach to land on
Mercury. The purpose of the payload is to measure the pressure, force, and acceleration of solar wind.
The measurements will be taken at thirty second intervals for forty-eight hours. To achieve functionality
at high temperatures, tourmaline piezoelectric sensors will be used. These sensors are functional up to
1000°C. If successful, Yoda will be caught on the solar winds by a solar sail and will measure these
factors as it is blown toward Mercury's surface. This means that Yoda will have to be forced closer to the
Sun and will have to travel backwards toward the surface of Mercury.
Table 1. Science Traceability Matrix
Science Objective
Measurement
Objective
Force
Determine the pressure of solar
winds from the sun to Mercury
Determine the acceleration of solar
winds from the sun to Mercury
Determine the force of solar winds
from the sun to Mercury
Measurement
Requirement
30 second intervals
for 48 hours
30 second intervals
for 48 hours
30 second intervals
for 48 hours
Acceleration
Force
Instrument Selected
Piezoelectric sensor
Piezoelectric sensor
Piezoelectric sensor
Table 2. Instrument Required Resources
Mass (kg)
0.0065kg
Power (W)
self-powered
Volume (cm3)
4.51cm3
Data Rate (bps)
continuous
Primary battery
0.09kg
Self-generating
N/A
---
UHF antenna
Processor
UHF transmitter
Total
0.0225kg
Negligible
Negligible
0.106kg
battery
battery
battery
---
N/A
Negligible
Negligible
4.51cm3
continuous
32bps
153,600bps
153,632bps
Instrument
Piezoelectric sensor
3.0 Alternative Concepts
The team separated into two groups and each group developed a design concept. The team developed two
concepts, evaluated both concepts, and used the strengths of each concept to develop a third concept that
would go on to become the final concept.
3.1 Concept 1: Concept 1 uses compressed helium method of deployment. The payload design features a
large balloon with a tethered Cube Sat directly below it. This Cube Sat is protected by four molded
Kapton plastic insulation and Nextel ceramic cloth heat shields. The inside of the balloon is doubled
layered, protecting five piezoelectric sensors. The data will be sent from the piezoelectric sensors to the
processor inside the Cube Sat, and finally, to the UHF transmitter, which will send it back to the
spacecraft.
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Payload Concept Proposal
Mercury Lander Mission
Spring 2012
3.2 Concept 2: In concept 2, the payload will be shot out of the spacecraft using compressed helium.
Yoda has a heat shield on one side with the Cube Sat directly behind it. To protect the Cube Sat from
possible tumbling and overheating the payload, spin stabilization will be used. Five piezoelectric sensors
that will be recessed into the heat shield will determine the force and acceleration. Data will be sent to the
internal instruments and, finally, to spacecraft.
3.3 Concept 3: Following the preliminary design review, the team evaluated concepts 1 and 2 and found
concept 2 to be the better concept. The team decided to move forward with the general design of the
shield and cube sat of concept 2 and to use the strengths of concept 1 to incorporate two solar sails into
the new design. In concept 3, Yoda is shot out using both the compressed helium deployment method, as
well as spin stabilization. The Cube Sat is protected by a single PICA heat shield with two Mylar solar
sails on the sides. The edges of the solar sails have grooves that fit in the grooves of the launcher. Two
tourmaline piezoelectric sensors are in each sail. Force and acceleration will be measured, processed, and
finally, transmitted to the spacecraft. This will continue for 48 hours, beginning one hour after
deployment.
Figure 1. Concept 1 Sketch
Figure 2. Concept 2 Sketch
Figure 3. Concept 3 Sketch
4.0 Decision Analysis
The team analyzed each concept using figures of merit. To help determine the best features of each
concept, each figure was given an overall weight depending on its importance. The first two concepts
were analyzed, and then the better of the two initial concepts were analyzed against a newer third concept.
4.1 Analysis of Concepts 1 and 2
The team decided that mass, power, safety, and stability were the most important aspects of the two initial
concepts. Neither of the two would greatly affect the instrumentation. The balloon was rated higher in
safety and aesthetics. The shield was rated higher in mass and stability. They both tied on their ratings of
power and ease of launch. After reviewing the requirements for the payload, the team decided that the
balloon was poorly suited for launch and drafted a new payload design combining the best aspects of both
designs.
Table 3. Payload Decision Analysis for Concepts 1 and 2
Figure of Merit
Weight
Mass
Power
Safety
Stability
Aesthetics
Ease of launch
Total
9
9
9
9
3
3
Concept 1
Balloon
3
9
9
3
9
3
252
Concept 2
Shield
9
9
3
9
3
3
288
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Reasons
Needs to fit mass constraint
Power will be the same
Instruments need to be safe
Payload must be free of tumbling
No reason
Must be launched without extra power
Payload Concept Proposal
Mercury Lander Mission
Spring 2012
4.2 Analysis of Concepts 2 and 3
After analyzing the first two concept sketches, another concept was developed to better meet the
requirements of the experiment. The same instrumentation was used, with the exception of a primary
400W hour battery instead of an Integrated Battery Daughter Board. The concept combined the best
aspects of the previous two sketches. Aesthetics and ease of launch were taken out of the new design
analysis because the team decided they were no longer important. After reviewing the new design and
calculating the FOM weights, the newest concept was more suitable.
Table 4. Payload Decision Analysis for Concepts 2 and 3
Figure of Merit
Weight
Mass
Power
Stability
Safety
9
9
9
3
Concept 2
Shield
9
3
3
3
Concept 2
Mickey-Mouse
3
9
9
9
Reasons
Needs to fit mass restraint
Power will be the same
Payload must be free of tumbling
Instruments need to be safe but of
less importance
Total
198
216
5.0 Payload Concept of Operations
Yoda will have two phases of operation: deployment and data collection. Prior to the spacecraft’s launch
from Earth, Yoda will be positioned inside an aluminum launch device. Yoda will be oriented so that it is
on the side of the spacecraft that will be closest to the sun on its approach to land on Mercury.
Phase 1: Deployment
 Launcher is pressured to 45.34 psi using compressed helium.
 The pressure will be released at a time and in a manner to cause Yoda to shoot out of the launcher
at a 90° angle to Mercury’s orbit.
 The grooves on the launcher will force Yoda to spin and prevent it from tumbling as it travels
towards the sun.
Phase 2: Data Collection
 The payload continues spinning with the heat-shield facing the sun and the Cube Sat on the
backside facing Mercury.
 Its life expectancy is approximately 7.35 days.
 Piezoelectric sensors placed in the solar sails will collect force and acceleration data for thirty
seconds and the transmitter will transmit data to the lander during the next thirty seconds.
 Yoda will continue the cycle of alternating measurements and transmitting every thirty seconds
for 48 hours.
 After recording these measurements, pressure can be determined by using p=F/A.
Figure 4. Payload Concept of Operations
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Payload Concept Proposal
Mercury Lander Mission
Spring 2012
6.0 Engineering Analysis
The F.O.R.C.E. found it necessary to perform analysis on several parameters that affect Yoda’s design.
The analyses include launch barrel analysis, launch pressure analysis, life expectancy analysis, and
battery mass analysis.
Launch Barrel Analysis.
First the team calculated the velocities of the 2cm and 4cm barrels using the pressure up to 4500psi. The
team graphed the data to compare the different barrel lengths. The team found that using the 4cm barrel
would be better because it used lower pressure to achieve the 100 m/s velocity needed to escape
Mercury’s sphere of influence.
Figure 5. Launch Barrel Analysis
Launch Pressure Analysis
Launch pressure was the next aspect of engineering analysis that was performed It was determined the
payload needed to have a final velocity of 100 m/s in order to escape Mercury’s sphere of influence so
that it can move toward the sun. Assuming constant pressure, the analysis reveals that a pressure of 45.34
psi will be required to properly launch Yoda.
Table 5. Launch Pressure Analysis
Calculations
vf2 = vi2 + 2ad
vf2 = 0 + 2 pAd/m
p = vf2m/2Ad
p = 45.34 psi
Substituted Equations
p = F/A
(Since F=ma, sub in for F)
p = ma/A
a = pA/m
p – pressure
m – mass
Variables
A – area (over which
pressure is applied)
a - acceleration
vf – final velocity (leaving
launch barrel)
vi – initial velocity (at rest)
F - force
d– length of barrel
Life Expectancy Analysis
The life expectancy of Yoda was the next engineering analysis that was performed. The team needed to be
certain that Yoda would survive long enough to complete the experiment. The team assumed the payload
would not survive beyond 0.2 AU, so the distance the payload will travel was determined to be the
difference between Mercury’s distance from the sun (0.4 AU) and 0.2 AU. The analysis revealed that
Yoda is expected to survive approximately 176.4 hours providing ample time to perform the experiment.
Table 6. Life Expectancy Analysis
Calculations
vf2 = vi2 + 2ad
vf = 47096.98646 m/s
Substitute vf in the
equation below.
Substituted Variables
a=GM- rv2/r2
d=(0.2AU)(1.49598 x 1011m)
vi = vsc + Δv = 4790 m/s + 100 m/s
t= d/vf
t = 635199.8769 s
t = 176.4444103 hr
Battery Mass Analysis
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Variables
vf – final velocity (at 0.2AU)
vsc – velocity of spacecraft
Vi – initial velocity upon launch
Δv – change in velocity to escape
Mercury’s sphere of influence
G – universal gravitational constant
6.67 x 10-11 N m2/ kg2
M – mass of sun = 1.9891 x 1030 kg
r – radius at .4 AU = 5.98392 x 1010 m
Payload Concept Proposal
Mercury Lander Mission
Spring 2012
The team decided it was necessary to determine the mass of primary batteries required. The team
calculated primary battery mass to be 0.09 kg. Initially, the team planned to use an integrated daughter
board with a mass of 2.2 kg as a power source. The battery mass analysis revealed that a 400 W Hr
primary battery with a mass of 0.09 kg would power Yoda’s instruments. Since the mass is lower, the
team decided to use the primary battery instead of the integrated daughter board.
Payload Structure Mass Analysis
The team decided it was necessary to calculate the mass of the heat shield and aluminum structure to
ensure that Yoda was within the 5 kg mass constraint. The team used the density of the materials (PICA
and aluminum) and the volume of materials required in the design to calculate the mass of the heat shield,
payload aluminum structure, and aluminum launcher. The team used the fact the density is an object’s
mass divided by its volume. With the density of the materials known and the volume of the structures
determines, it was a matter of solving the equation D = m/v for mass. The team calculated the mass of the
heat shield to be 1.296 kg, the mass of the aluminum launcher to be 0.57 kg, and the mass of the
aluminum structure to be 1.1 kg.
7.0 Final Design
The final design features a PICA heat shield and two Mylar covered canisters holding the piezoelectric
sensors. The Cube Sat is located directly behind the PICA heat shield, and the two canisters are connected
to it from the sides. The dimensions of the payload are 42cm x 10cm x 18 cm. The launcher contains
grooves to better fit the payload and send it into spin stabilization upon launch with the aid compressed
helium.
Figure 6. Front view of payload
Figure 7. Back view of payload
Figure 8. Launcher
Table 7 Final Concept Mass and Power Budget
Component
Number
Mass (kg)
Power (W)
Self-powered
Total Mass
(kg)
0.026kg
Total Power
(W)
Self-powered
Piezoelectric
Sensor
Primary
Battery
UHF Antenna
4
0.0065kg
1
0.09kg
Self-powered
0.09kg
Self-powered
1
0.0225kg
Battery
0.0225kg
Battery
Cube Sat (2U)
Processor
Heat shield
Transmitter
Solar Sails
Aluminum
Total
1
1
1
1
2
1
---
1.33kg
Negligible
1.296kg
Negligible
0.0215kg
1.67kg
---
0W
Battery
0W
Battery
0W
0W
---
1.33kg
Negligible
1.296kg
Negligible
0.043kg
1.67kg
4.4kg
0W
Battery
0W
Battery
0W
0W
---
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