ENAE 483: Principles of Space System Design Loads, Structures, and Mechanisms

Transcription

1 ENAE 483: Principles of Space System Design Loads, Structures, and Mechanisms Team: Vera Klimchenko Kevin Lee Kenneth Murphy Brendan Smyth October 29 th, 2012

2 Presentation Overview Project Overview Mission Guidelines Project Specifications Initial Design Choice Crew Capsule Assumptions for Model Launch Acceleration Loads Random Vibration Thermal Stress Pressurization Launch Loads Summary Buckling Force Docking Loads Earth EDL Splashdown Landing Vehicle Landing Vehicle Sizing Sizing of Propellant Aestus Rocket Engine Description/Analysis of Landing Gear Buckling Analysis Shear Stress Analysis

3 Mission Guidelines Calculate loads and conduct structural analysis for crewed spacecraft and design lunar landing vehicle Crew of three for 10-day nominal mission (plus three contingency days) We have already designed the crew life-support, power, propulsion, and thermal systems in previous design reports

4 Project Specifications Structure of the crew cabin supports all significant loads sources during: Earth launch Pressurization loads Docking loads Lunar landing loads Earth EDL

5 Project Specifications (cont.) Basic design of landing vehicle in terms of tank size, engine, etc. Inert mass 2199 kg Propellant mass (N2O4/MMH) 9914 kg Landed payload includes crew vehicle and fully fueled propulsion stage (gross mass 12,110 kg) Design and analysis of the landing gear Touchdown velocity 3 m/sec vertical, 1.5 m/sec horizontal Includes impact attenuation to prevent bounce Includes design of deployment articulation and actuators

6 Choice of Initial Design We chose to use the design of Team B4 of the Power Propulsion and Thermal Project. Their last report was fairly comprehensive with propulsion and moment of inertia calculations which we thought would help when trying to do analysis of design

7 Crew Vehicle: Material Values Acceleration due to Gravity g = 9.8 m/s 2 For Aluminum: Young s Modulus ( E = 69 gigapascals ) Yield Stress ( Y s = 256 megapascals ) Poisson s Ratio ( ν = 0.34 )

8 Relevant Parameters P cabin = 59.7 kpa t wall = 10 cm α = 25 H = 3.83 m α t = m m C

9 Modeling the Cone Radius R is the maximum at the bottom of the cone and is a function of y and the half-angle: R = (H y)tan (α) I is given by π 4 R o 4 R i 4 however assuming thinwalled cone allows us to approximate this as πr 3 t wall This cone is unlike a cylinder in that the radius is constantly changing in order to approximate, for the rest of our calculations we assume: I = πrt wall

10 Modeling the Cone (cont.) Assuming a hollow uniform cone the center of gravity is given by: h cg = H/4 Area of the cross section of the cone is given by: A = 2πRt wall

11 Modeling the Cone (cont.) For stress calculations due to acceleration and vibration, our cabin was approximated by a bottom mass on the bottom of the craft with a cone on top The bottom mass accounts for the fact that much of the cabin mass lies within the bottom of the vehicle because propellant tanks and batteries were placed there, and most of the life-support systems lie on the bottom of the cabin at a height of l W bottom = 2000 kg l bottom = 0.5 m

12 Loads due to Launch Accelerations Assuming the following values for launch acceleration (y is up with base on ground): g y = 9 g g x = 5 g g z = 5 g g trans = g x 2 + g z 2 = 7.07g

13 Launch Accelerations (cont.) Assuming y-axis acceleration is absorbed only by the bottom mass, stress is given by σ LA = MR I + W bottom A g y M = g trans (W cone h cg + W bottom h bottom ) Plugging in our parameters M = Nm σ LA = Pa

14 Random Vibrations Frequency (Hz) PSD (g 2 /Hz) f n ζ <150 Hz Hz.020 >300 Hz.005 From Akin

15 Random Vibrations (cont.) f 1 = π EIg W bottom L 3 +W cone L 3 = 207 Hz RLF 1 = πf 1 PSD 4ζ = 9.01 g M = Nm σ vib = Pa

16 Thermal Stress Length of the wall of cone given by s = L/cos (α) Assuming a change of 38 C δs = α δt l =.0018 m Assuming the support structure changes only half as much as the side of cone Thermal stress is then σ thermal = E δs s 0. 5 = Pa

17 Pressurization Loads Meridional Stress at max radius given by σ pm = P cabinr 2t wall cos α Hoop Stress is given by σ pm = P cabinr t wall cos α = Pa = Pa

18 Earth Launch Totals Load Source Limit Stresses (Pa) FOS Design Stress (Pa) Launch Acceleration Pressurization Random Vibration Thermal Total MS = Y s σ stress 1 = = 431% Our crew cabin is significantly overdesigned!

19 Buckling of Cone Buckling of a normal thin-walled cone can be estimated by P cr = 2πEt 2 wall 3 1 ν 2 cos α 2 = N This is the maximum axial force that the cone can experience before buckling See References: Simple solutions for buckling of orthotropic conical shells

20 Capsule Docking Rates Parameters from PPT project: t impulse = 0.06 seconds T roll = 50.1 newtons T yaw = T pitch = newtons r arm = 1.18 meters

21 NASA Limits for Capsule Docking Rates From the International Docking System Standard (IDSS) Interface Definition Document (IDD)

22 Spin Rate: α = τ I zz Capsule Docking Rates (cont.) = 2 T roll r arm I zz θ = α t impulse 180 = deg/s π θ is within limit set by NASA

23 Capsule Docking Rates (cont.) Yaw/Pitch Rates: Due to angled thruster: T total = T yaw + [T yaw sin 65] α = T total r arm I yy θ = α t impulse = 0.12 deg/s θ resultant = θ 2 +θ 2 = 0.17 deg/s θ resultant exceeds NASA limit by 0.02 deg/s Can be resolved by using a faster solenoid

24 Capsule Docking Rates (cont.) Closing (Axial) Rate: a = 2 T yaw m total v = a t impulse = m/s v is within limit set by NASA Lateral Rate: a = 2 T roll m total v = a t impulse = m/s v is within limit set by NASA

25 Capsule Docking Load Kinectic Energy = m v2 2 = k x2 2 k = m v2 ; x = 0.4m x 2 Stiffness and force values calculated for each docking rate Total load on docking hatch: F total A dock σ docking = 12.7 kpa

26 Earth EDL Splashdown Assumptions: Capsule is travelling at ~25mph (11.18m/s) at splashdown Elapsed time of deceleration during splashdown is approximately 2 seconds F = m capsule v t F = 26795N (in the axial direction) Less than the critical buckling load σ splashdown = F A base = 2.7kPa

27 Landing Vehicle Sizing The landing vehicle: Oxidizer: N2O4 Fuel: MMH Aestus Rocket Engine Deployable landing gear Oxidizer and Fuel Information: Oxidizer to Fuel Ratio 2.05 Density of N2O4 Density of MMH Total Mass of Propellant Volume of N2O4 Tank Volume of MMH Tank Mass of MMH 1400 kg/m^3 868 kg/m^ kg 3251 kg Mass of N2O kg 4.76 m^ m^3

28 Sizing of the Propellant (cont.) If spherical tanks were considered for the fuel and the oxidizer. diameter of diameter of diameter of N2O4 tank 3.06 m MMH tank 2.82 m the landing vehicle 5.9 m If both of the tanks were spherical, the diameter of the landing vehicle would need to be equal or greater than 5.9 meters. This is unacceptable because it is 1.6 times larger than the intended diameter of the landing vehicle whish is 3.57 meters. The tanks will be cylindrical with hemispherical end caps. Height of N2OH Tank=2.05 meters Diameter of N2OH Tank= 1.46 meters Height of MMH Tank=2.05 meters Diameter of MMH Tank=1.28 meters

29 Height (m) Sizing of the Propellant Tanks 10 Sizing the N2O4 and MMH Tanks Height (m) vs. Diameter of the Hemispherical Ends(m) N2O4 MMH Diameter (m)

30 Aestus Rocket Engine Oxidizer/Fuel N2O4/MMH Average Thrust 27.5kN Mixture Ratio 2.05:1 Isp 324 sec Nozzle Area Ratio 83:1 Gimbal Capability +/- 6 degrees Engine Mass 136 kg

31 Description of the Landing Gear The landing gear will consist of 4 retractable legs. Each leg a length of 2.33 meters, extends at an angle of 45 degrees from the spacecraft. The weight of the landing vehicle and the crew vehicle will be distributed evenly between two legs. All of the legs will have a honeycomb cartridge that will absorb some of the touchdown energy. The honeycomb mesh displacement is meters

32 Landing Gear The landing gear will have three joints. Joint #1: Joint one will enable the leg to rotate down and stop when the leg is parallel to the body. Joint#2: Joint two will enable to lower part of the strut to bend underneath the landing vehicle. Joint #3: Joint three will enable the foot pad to rotate under the landing vehicle. 4 Legs with 90 degree separation Stowed

33 Landing Gear Analysis Vx 1.5m / s Vy 3m / s V m / s Mtot Mleg 3,833kg 4 4 ( M )( V ) k 65,112N / m 2 d F max k * d 52,900 Spring Constant Fmax Fmax d=max displacement for the honeycomb mesh

34 Landing Gear Specifications Choice Choosing the angle between the vehicle body and the fully deployed leg: 45 degrees If the angle is too shallow the vehicle could encounter stability issues and over turn. If the angle is too low the engine nozzle will hit the ground upon touchdown. Choosing the length of the strut: 2.33 meters The length of the strut was chosen so that there would be enough clearance between the nozzle and the surface of the moon. Choosing the geometry of the strut: thin-walled hollow tube Ri=.14meters Ro=.10meters tube was chosen to withstand large buckling loads and shear loads. The inner and outer radius of thin wall tube were chosen after doing trade studies presented below.

35 Landing Vehicle Sketch

36 Buckling and Shear Stress Analysis Fstrut 53,000 Fstrut Fstrut SF 1.5 Fbuckle 2 EI Fbuckle 2 ( KL) 4 I ( r0 rin 4 L Length of E I Youngs Modulus Fbuckle ) Area Moment of 4 the Primary Strut Inertia F F sin( ) A A Yield Stress of Aluminum

37 Buckling Force (N) Buckling Analysis Buckling Force (Newtons) vs. Area Moment of Inertia (m^4) 6E+09 5E+09 4E+09 3E+09 2E+09 1E Area Moment of Inertia (m^4)

38 Area Moment of Inertia (m^4) Buckling Analysis (cont.) Area Moment of Inertia (m^4) vs. Outer Radius(m) Outer Radius (m) **The force due to buckle has a dependency on the moment of inertia. Moment of inertia is dependent on the outer and inner radius. The inner radius was varied constantly with regard to the outer radius.

39 Mass Ikg) Buckling Analysis (cont.) Mass (kg) of the Primary Strut vs. Radius (m) Mass Individual Mass Total Radius (m)

40 Shear Stress (MPa) Shear Stress Analysis Shear Stress (MPa) vs. Radius (m) Radius (m)

41 References International Docking System Standard (IDSS) Interface Definition Document (IDD). Publication. NASA, 13 May Web. 28 Nov < pdf>. Tong, Liyong, et al. Simple solutions for buckling of orthotropic conical shells. International Journal of Solids and Structures: Volume 29.8, 1992, Pages