Loads, Structures, and Mechanisms Design Project Team C4 Jason
Burr, Rebecca Foust, Samantha Johnson, Kiran Patel, and Dennis
Sanchez
Slide 2
Mission Objectives Structural Analysis: Crew Vehicle Lunar
Landing Vehicle Crew Vehicle Earth Launch Pressurization Loads
Docking Loads Lunar Landing Loads Earth EDL Landing Vehicle Basic
design Inert mass: 2199 kg Propellant Mass (N2O4/MMH): 9914 kg
Payload: 12,110 kg Landing Gear Analysis Touchdown velocity 3 m/s
vertical, 1.5 m/s horizontal
Slide 3
Crew Capsule Selection Jasons crew capsule was selected because
it has no external elements like radiators or solar arrays It also
has the highest mass margin, so we have the most mass available for
our structural design
Slide 4
CREW VEHICLE
Slide 5
Pressurization Stresses in a Conical Head Analogous to the
maximum stresses in the cylinder, there are maximum stresses in the
conical headHowever, in contrast to the cylinder, it is not
possible to establish simple expressions for the three stress
indexes Stresses in a Pressure Vessel With a Conical Head For this
reason we will model our cone as a cylinder to find pressurization
loads.
Slide 6
Pressurization Loads Cabin Pressure throughout mission =
59.77kPa Hoop stress on cabin = 1.067 MPa Longitudinal stress on
cabin = 533 kPa Hoop stress will cause the structure to fail before
longitudinal stress can cause failure Total Pressurization Stress =
1.067 MPa Pressurization stress will occur at all times and is
added into all total stress values
Slide 7
Earth Launch Use Falcon Heavy to launch to LEO Can carry 53,000
kg to LEO Max Thrust = 11,200 kN Force at payload = 587 kN Stress
from thrust = 538 kPa Random Vibrational Stress = 181 kPa
Slide 8
Temperature Variation in Atmosphere
Slide 9
Earth Launch Thermal loading occurs as payload travels through
different layers of the atmosphere. Max temperature difference = 80
K Thermal stress = 122.5 MPa Total Stress at Launch = 124.3
MPa
Slide 10
Lunar Landing Loads Maximum G force that our capsule will
undergo is 1.125g This is determined from the maximum G force the
astronauts can undergo while standing Force from Lunar Landing =
186 kN Total Lunar Stress = 1.24 MPa
Slide 11
International Docking System Standard (IDSS) Maximum force
exerted during seal closure Total Docking Stress = 1.16 MPa
Slide 12
Earth EDL - Heat Shielding AVCOAT ablative heat shield Total
Thermal Loading: 2597.9 MPa Used on Apollo crew modules Will
diffuse heat into the air as opposed to the structure
Slide 13
AVCOAT Shielding Will withstand total thermal loading for EDL
Epoxy resin in fiberglass honeycomb matrix Above before EDL Below
after EDL
Slide 14
Earth EDL Maximum G force that our capsule will undergo is
7.19g Force during EDL = 493 kN Thermal Stress = 122.5 MPa This is
from fluctuations in the atmosphere. The heat shield takes all of
the thermal stress during EDL Total Stress during EDL = 126
MPa
Slide 15
Safety Factors NASA Technical Standard: STRUCTURAL DESIGN AND
TEST FACTORS OF SAFETY FOR SPACEFLIGHT HARDWARE Need a safety
factor of at least 1.25 to fall under NASA standards
Slide 16
Highest Stress on Lander During EDL the stress reaches 126 MPa
Based off of the yield stress of different materials, aluminum will
be the best material for our structure With a safety factor of 1.25
the stress is still well under the yield strength of aluminum, 386
MPa
Slide 17
Yield Stress vs. Density
Slide 18
Total Stresses CaseDesign Stress with SF (MPa) Pressure1.33
Launch155.3 Lunar Land1.55 Docking1.45 Earth EDL157.5 MOS145.0% MOS
is based off of case of maximum load
Slide 19
LANDING VEHICLE
Slide 20
Lander Strut Configurations Three Distinct Design Possibilities
1.Rigid Structure 2.Spring and Damper Attenuation 3.One-Time Energy
Dissipation
Slide 21
Lander Strut - Rigid Structure Advantages Simple analysis
Re-usable Easy deployment Disadvantages Large loads Crash landing
scenario Large magnitude accelerations Potentially fatal to
astronauts
Slide 22
Lander Strut Spring and Damper Advantages Re-usable Adjustable
maximum accelerations Disadvantages Complicated analysis
Challenging deployment technique Springs act to move the struts to
their equilibrium positions
Slide 23
Lander Strut Energy Dissipation Advantages Relatively simple
analysis Easy deployment Comparable to rigid strut Low
accelerations Disadvantages One-time use
Slide 24
Design Choice Energy Dissipation Honeycomb Energy Dissipation
Wide range of strengths available Constant force during crushing
Reliable energy dissipation
Slide 25
Honeycomb Energy Dissipation Energy Conservation Kinetic Energy
to Crushing Work L crush is the total length of Honeycomb required
to dissipate all of the energy from landing Acceleration increases
with decreasing stroke length
Slide 26
Honeycomb Acceleration Limits Maximum Acceleration 1.125g Limit
imposed on elevators Provides low enough acceleration that
astronauts can remain standing in lunar descent Maximum Stroke
Length 1.75 m Limit to easily store struts in descent stage Assume
all energy is dissipated in a single strut (worst case) Results in
minimum acceleration of 0.328g
Slide 27
Honeycomb Strength Selection Various Honeycombs of different
strengths and densities Performing the worst case landing, we can
determine the Honeycomb mass and diameter to meet our
constraints
Slide 28
Strength Selection Continued Optimum Honeycomb: p=8.1 lb/ft 3 P
crush =750psi Note: All designs from this point on use this
Honeycomb
Slide 29
Honeycomb Landing Scenarios Three likely scenarios to arise
are: 1.Landing on a single strut Landing on extremely uneven
surfaces (rock) 2.Landing on a constant incline Assumed smooth
planar surface 3.Landing on a flat surface Specific type of
incline
Slide 30
Case 1: Uneven Landing
Slide 31
Landing Scenario Single Strut Worst case all energy dissipation
is in a single strut As before, the length of the crushed section
is:
Slide 32
Case 2: Sloped Landing
Slide 33
Landing Scenario Sloped Landing More complicated Total length
required for energy dissipation remains the same Maximum crushing
occurs in the leading strut of the lander, minimum in the trailing
one Assumptions: Lander remains horizontal to the surface during
descent Maximum landing slope is determined where the trailing leg
does not need to absorb any additional energy
Slide 34
Sloped Landing First Pass Note: assume the lander must make
contact with all four struts any other configuration is unstable!
(This is the cause of the curves ending) *Maximum crushing occurs
in the leading strut. It never reaches the maximum crush length
because the two struts between the leading and trailing one absorb
some of the energy.
Slide 35
Sloped Landing Refined Pass Considering only the data points
that maximize the slope at any given acceleration, we can produce
the following plot: Possible landing slope is maximized at the
lowest acceleration, 8.07 and 0.328gs, respectively
Slide 36
Case 3: Zero-Slope Landing
Slide 37
Landing Scenario Flat Surface Subset of the previous scenario
where the slope is equal to 0 Energy is dissipated evenly between
the four struts thus the crush length is as well
Slide 38
Summary of Landing Crush Lengths StrutSingle Strut LandingMax.
Slope Landing (8.07)Min. Slope Landing (0) Leading Strut1.75 m1.56
m0.438 m Mid-Strut0 m0.970 m0.438 m Mid-Strut0 m0.970 m0.438 m
Trailing Strut0 m 0.438 m The following table represents the worst
crush lengths for the three main landing scenarios: *Honeycomb
Mass= 2.62 kg/strut Where: a = 0.378gsP crush = 5.17 kPa A crush =
0.011 m 2 m = 16905 kgL crush = 1.75 m crush = 130 kg/m 3
Slide 39
Landing Strut Analysis Source of loads on the landing struts:
1.Earth launch loads 2.Lunar landing loads 3.Thermal loads Landing
struts are 6 m long before crushing Minimum of 4.25 m after
crushing Model the struts as hollow tubes Design varied to
minimized margin of safety Neglect joint forces
Slide 40
Strut Analysis Earth Launch Stress due to launch forces and
moments Iteratively solved to minimized mass with: ax = 8.5gay =
5.8gaz = 4.85g L = 6 m
Slide 41
Strut Analysis Lunar Landing Stress due to landing force and
moment Assumed the landing force is purely axial and purely
rotational Physically impossible to occur at once, but creates an
extreme-upper bound on loading Iteratively solved to minimized mass
with: a = 0.328g L = 6 m
Slide 42
Strut Analysis - Thermal Rapidly changing temperatures while
during Earth launch Greatest temperature variation ~80 K
Iteratively solved to minimized mass with: T = 80 K L = 6 m
Slide 43
Strut Analysis Combined Loading Consider Earth and thermal
loading combined, as well as lunar landing and thermal loading All
cases use factors of safety (SF) of 1.4 Iteratively design with
various radii to minimize mass and the margin of safety or
Depending on which is the limiting (lower) value
Slide 44
Strut Analysis Analyzed Materials MaterialE (GPa)p (g/cm 3 )
(m/m*K) (MPa) Aluminum 2024722.7822.2324 Aluminum 7075712.7822.2490
Titanium Ti-6Al-4V1104.468.6869 Steel AISI43402007.8131483 Steel
300M2007.8131520 Consider the following metals for our struts:
These materials are used with the preceding formulas to find the
optimum strut design that 1.Minimizes strut mass 2.Keeps an
external strut diameter below 20 cm 3.Has a MoS = 0
Slide 45
Strut Analysis Optimization Titanium appears to be the best
metal, but
Slide 46
Strut Analysis Optimization Cont. When we consider the
constraint of less than 20 cm (external) diameter, we see that
Steel 300M at this point is 68.24 kg/strut, whereas titanium at an
external diameter of 20 cm is 72.78 kg/strut
Slide 47
Strut Analysis Summary Landing Strut Total length 6 m Outer
radius10 cm Inner radius9.5 cm MaterialSteel 300M MoS0 Strut
Mass68.24 kg/strut Honeycomb Max crush length1.75 m Min crush
length0 m Honeycomb Mass2.62 kg/strut
Slide 48
Strut Storage and Articulation The landing struts are divided
into three 2-meter segments: two hollow tubes and one honeycomb
piston During launch, the three segments are stored side by side as
shown here Rotary actuators at the joints align the segments and
lock them into place for lunar descent
Slide 49
Actuators The landing struts will be stored folded into 3
sections during launch Before lunar descent, the landing struts
will deploy using the space-rated rotary actuator shown below This
actuator was chosen because it can produce high torque at low
speeds. Lower speeds will reduce vibrational loads on the
spacecraft
Slide 50
Supplemental Ladder The honeycomb piston is designed for
worst-case scenario loading, so a best-case load would result in
significantly less compression Because of this, a supplemental
ladder is attached to the end of the second segment on the ladder
strut The ladder will deploy in all scenarios, but will only be
critical for light load cases where the astronauts would otherwise
be unable to perform EVAs
Slide 51
Footpad Design The footpad needs to be large enough to remain
stable across small variations in surface depth We chose a radius
of 0.5 m because this covers a total area of 0.7854 m 2, which
should be sufficient to ameliorate the effects of an undesirable
landing site The footpads are mounted with a ball joint to allow
them to rotate as needed for various surface slopes
Slide 52
PROPULSION MODULE
Slide 53
Engine Mount Structure We are using a truss design to mount the
engine to the module Needs to handle a maximum thrust load of 15568
N from engine The design is composed of 8 criss-crossing trusses
which distribute the weight evenly amongst them Top View of Truss
Engine Mount Engine
Slide 54
Side View of Truss Engine Mount Engine Truss Mount
Slide 55
Truss Analysis In the analysis the truss was scaled by.2 and
the force applied at each joint was (1/8)(Maximum Thrust)/10 or 195
N
Slide 56
Truss Design Each truss member is a hollow tube composed of
aluminum and carries a maximum load of about 4070 N The radius of
each member is 5 cm The thickness of each member is chosen to be 1
cm Maximum Stress = 1.43 MPa Well under the yield stress of
aluminum, 386 MPa
Slide 57
Truss Member Stress vs. Thickness
Slide 58
Propulsion Module Requirements The full propulsion module is a
two meter long cylinder with a diameter of 3.57 m Four 60x30 cm
sections are cut from the overall cylinder to store the landing
struts Required propellant volumes: This volume also holds the
engine and engine mount, which occupy a total volume of 0.713 m 3
Monomethyl Hydrazine (MMH)Nitrogen Tetroxide (N 2 O 4 ): 4.333 m 3
4.208 m 3
Slide 59
Slide 60
Tank Sizing The tanks were cylindrical with ellipsoidal end
caps. The height of the ellipse was modeled as 0.25*radius of
cylinder Using this equation and the volumes stated earlier
produces the following radii, which fit well within our design
limits: Monomethyl Hydrazine (MMH)Nitrogen Tetroxide (N 2 O 4 ):
0.4227 m0.4165 m
Slide 61
Mass Totals ComponentMass (Kg) Crew Systems1500 Power,
Propulsion, Thermal4795 Aluminum Shell1687 Propellant9914
Propulsion Inert Mass2199 Landing Struts283.4 Total Mass20,378
Slide 62
Final Design
Slide 63
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