Loads, Structures, and Mechanisms Matthew Marcus Chris O'Hare Alex Slafkosky Scott Wingate
Presentation Overview Design requirements Initial crew capsule design choice Pressure vessel design Pressure loads Earth EDL loads Thrust structure design Landing gear design Landing orientation Strut design Impact attenuation 3D Rendering Mass estimations References 2
Design Requirements The structures of the lunar lander must be designed to handle the following loads: Earth launch Crew cabin pressurization Orbital docking Lunar landing Earth EDL Three structures will accomplish this Pressure vessel Thrust structure Landing gear 3
CS and PPT Design Choice I Thruster Quads Water Tanks Radiator Panels Fuel Cell Fuel Tanks Propulsion Fuel Tanks Fig. 1 Group B4's capsule design from PPT project 4
CS and PPT Design Choice II This design was selected because its radiators sit flush on the capsule surface and uses fuel cells instead of deployable solar panels The lack of such deployable panels means less structure is required for support during launch and landing This reduces the required overall structural mass, allowing for more payload 5
Crew Capsule: Assumptions For purposes of loading analysis, the crew capsule is assumed to be the maximum mass allowed, 4795 kg All systems internal to the crew capsule are assumed to be designed to withstand mission maneuvers Therefore, this analysis will focus on designing the pressure vessel to withstand all loads with an added safety factor of SF = 1.3 6
Pressure Vessel: Load Analysis Four loading conditions: Crew cabin pressurization Earth reentry deceleration Parachute deployment Orbital docking maneuvers Use largest loading condition to choose pressure vessel skin thickness Pressure vessel material: 6061-T4 aluminum Yield Strength: 145 MPa Modulus of Elasticity: 69 GPa Density: 2.7 g/cc 7
Pressure Vessel: Pressurization Loads Cabin is pressurized to 69 kpa Using cylindrical approximations for pressure load analysis, the minimum thickness is determined by the hoop stress at the point of max radius: tmin=1 mm Fig. 2 8
Pressure Vessel: Atmospheric Reentry Loads Maximum allowed deceleration is 7g assuming a reentry profile similar to the Apollo missions Modeling the aerodynamic load as a compressive force, the pressure vessel thickness must prevent buckling: tmin=1.65x10-4 mm 9
Pressure Vessel: Parachute Deployment Loads Deceleration by parachute deployment shall not exceed 4g Parachutes deploy from top of capsule, exerting a tensile stress on the pressure vessel: tmin=0.7 mm 10
Pressure Vessel: Docking Loads Orbital docking is modeled as an impulsive compressive force Because docking maneuvers are conducted at much lower acceleration than the earth reentry case of 7g, the docking force will be much less than the aerodynamic load Therefore, a pressure vessel designed to withstand reentry will also withstand docking 11
Pressure Vessel: Stress Concentrations Discontinuities in the pressure vessel for the window and door introduce stress concentrations around the opening. Stress concentration factor (CF) is conservatively approximated to be 10 in lieu of physical tests Therefore the minimum thickness is defined by the pressure stress concentrations: tmin=10 mm m=795 kg 12
Lunar Lander Design Design Considerations The thrust structure must be able to withstand all accelerations in every mission scenario The landing structure must be able to withstand all landing forces in all considered configurations The landing structure must decelerate the vehicle in such a way as to maintain the safety of the crew 13
Thrust Structure Load structure modeled as simple triangular truss: Marmon Ring Fig. 3 14
Thrust Structure: Max Launch Acceleration Max acceleration = 4.4g FAB = 96 kn FAC = 324 kn FBC = 100 kn Fig. 4 15
Thrust Structure: Max Lunar Descent Acceleration Max acceleration = 0.28g FAB = 13 kn FAC = 13 kn FBC = 8 kn Fig. 5 16
Thrust Structure: Resting on Lunar Surface Max acceleration = g /6 FAB = 50 kn FAC = -56 kn FBC = -43 kn Reaction Force from landing leg Fig. 6 17
Considerations Velocity Maximum landing velocity Landing Orientation Design for full braking force capable of being taken on one leg worst case scenario Considering both spacecraft orientation and landing site slope, assume a combined maximum landing offset of 12 from upright (Apollo legacy) Sliding friction Friction on surface of moon = 0.08 Small enough to ignore landing stress due to friction in structural analysis 18
Orientation I Side view of landing orientations: Fig. 7 19
Orientation II Top-down view of landing orientations: Fig. 8 20
Considerations I At least one crew member is in a standing position for descent control Assume some restraint harnessing in both spacecraft local vertical and local horizontal Landing force must be small enough such that the pilot maintains continuous control of the craft Maintain grip on control system Maintain position in front of control system 21
Considerations II These are the maximum accelerations in local spacecraft during landing, to be used to drive landing gear design: y-axis: 1.0 g x-axis: 0.20 g 22
Strut Orientation I As we are ignoring friction with the surface, initial impact velocity is only considered along the vector normal to the surface The landing strut angle is optimized such that the maximum impact velocity travels along the strut This minimizes transverse velocity, and thus minimizes bending moments along the landing strut and the whole landing system Allows for lighter landing struts to be used 23
Strut Orientation II Fig. 9 Fig. 10 Want to minimize Vtrans 24
Strut Orientation III Figure 11 shows the resultant longitudinal strut impact velocity at lunar surface angles. Each line represents a strut angle offset from spacecraft local vertical Figure 12 shows the resultant transverse strut impact velocity at lunar surface angles. Each line represents a strut angle offset from spacecraft local vertical 25
Strut Orientation IV Fig. 11 26
Strut Orientation V Fig. 12 27
Strut Orientation VI The chosen strut orientation had the smallest maximum transverse impact velocity across all landing orientations Strut angle offset = 6 Maxima Transverse impact velocity = 0.339 m/s Longitudinal impact velocity= 3.23 m/s 28
Strut Orientation VII Fig. 13 29
Strut Orientation VIII Fig. 14 30
Tipping In order to prevent tipping due to transverse forces, the landing gear feet must be placed wide enough to prevent this. yc = 0.74 m X=2.1 m Fig. 14 31
Strut Sizing I Truss struts are designed to withstand the largest expected stresses All final sizes assume a safety factor of 1.4 We will be using Al 6061-T6 Aluminum for truss construction E = 69 GPa σyield= 275 MPa 32
Strut Sizing II Axial Force and Buckle Considerations: Strut Amin (cm2) Fmax (N) Pbuckle f(i) (N) Ideal SF Ideal SF Ideal SF A 1.41 e5 1.97 e5 5.13 7.18 1.39 e11*i 9.93 e10*i B 1.41 e5 1.97 e5 5.13 7.18 5.98 e10*i 4.27 e10*i C 4.75 e4 6.65 e5 1.72 2.41 2.97 e12*i 2.12 e12*i Moment Considerations: Strut B Mmax (N-m) y/i Ideal SF Ideal SF 8.08 e4 1.13 e5 3403 2433 33
Strut Sizing III The primary design factor for the longitudinal strut is the transverse bending load Fig. 15 34
Strut Sizing IV Main Strut Side Strut Radius (m) 12.2 2 Thickness (cm) 1 1 I (m4) 5.04 e-5 2.70 e-5 Mass (kg) 59.74 1.22 Mass Total Strut (kg) 60.96 Misc Support (kg) 10 Total per leg (kg) 70.96 Structure Total (kg) 283.84 35
Strut Sizing V The landing structure will consist of four landing legs, each of which employs a truss design Each leg truss consists of two main members Main strut (A,B) a continuous piece that experiences the calculated longitudinal forces Side strut (C) piece that experiences the calculated transverse forces. It is attached to the main strut via a hinged sleeve for stowing configuration purposes 36
Strut Sizing VI The forces are transferred from the landing struts into the propulsion module structure Braking forces are absorbed by honeycomb structures on both the main and side struts Fig. 16 37
Honeycomb I Objectives: Attenuate shock of landing on the lunar surface Ensure system does not "hop" after initial touchdown Modal: Acceleration constraints Optimized strut angle Energy Dissipation Equations 38
Honeycomb II Acceleration constraints: al=0.851 g at=0.286 g Strut angle optimization yields possible velocity profiles at different surface angles: VL VT 39
Honeycomb III Energy Dissipation: Where: m = mass fcr= Crush Pressure η = Safety Margin Sprecrush= Precrush Length 40
Honeycomb IV Fig. 17 41
Honeycomb V Fig. 18 42
Honeycomb VI Fig. 19 43
Honeycomb VII The required honeycomb dimensions are designed to handle the worst-case scenario velocity in both the longitudinal and transverse directions Longitudinal Transverse Length (cm) 89.8 3.57 Area (cm2) 32.9 11.1 44
3D Rendering: Lunar Propulsion Module 45
3D Rendering: Lunar Propulsion Module 46
3D Rendering: Full Lander Assemble 47
Mass Summary System Mass (Kg) Crew Systems 3219 Pressure Vessel 795 Thrust Structure 153 Landing Gear 284 Lunar Ascent Stage 12110 Total 48 16561
References Elverum, G., P. Staudhammer, J, Miller, A Hoffman, and R. Rockow. "The Descent Engine for the Lunar Module." Proc. of AIAA 3rd Propulsion Joint Specialist Conference, Washington D.C. Web. Hexcel Corporation. HexWeb Honeycomb Energy Absorption Systems: Design Data. Stanford: Hexcel Corporation, 2005. Web. Hexcel Corporation. HexWeb Rigicell: Corrosion Resistant Aluminum Corrugated Honeycomb, Product Data. Stanford: Hexcel Corporation, 2006. Web. Rogers, William F. Apollo Experience Report - Lunar Module Landing Gear Subsystem. Rep. N.p.: n.p., 1972. NASA TN D-6850. Scribd. Web. 28 Nov. 2012. <http://www.scribd.com/doc/46962760/apollo-experience-report-lunar-module-landing-gearsubsystem>. Rogers, William F. Apollo Lunar Module Landing Gear. Rep. NASA, n.d. Web. 49