Avionics, Software, and Simulation ENAE483 Fall 2012

Transcription

1 Avionics, Software, and Simulation ENAE483 Fall 2012 Team D7: Michael Cunningham Matthew Rich Michelle Sultzman Scott Wingate

2 Presentation Overview Project Specifications Crew Capsule Design Choice Communications Link Budgets o Earth Communications o EVA Communications Antenna Design/Placement Sensor Specifications o Capsule Atmosphere o Dynamics o Miscellaneous Design/Build/Test Concepts References 1 D7

3 Project Specifications Avionics system design for a 13 day, 3 man mission to the moon Design communications systems to give positive link margins for communications with earth and surface EVA operations List capsule sensors, their criticality, and their sampling frequency Develop ideas for human factors tests that can be performed to validate capsule design D7 2

4 Crew Capsule Design Choice D7 3

5 Crew Capsule Design Choice This crew capsule design was chosen due to the size of the hatch, which would facilitate ingress and egress operations The folding chairs allow more room in the capsule for the three astronauts to put on spacesuits The capsule's three windows give added visibility during critical mission maneuvers There is empty space in the nose cone for avionics equipment D7 4

6 Link Budget Analysis: Assumptions The following link budget analysis makes the following standard assumptions: o Total System Loss = o Bit Error Rate = 1*10-5 o Data Rate = 2.8*10 4 bits/sec o Eb/No Required = 9.4 Target link margin 3 db o Determine how much transmitter power is needed to yield at least 3 db of link margin for each connection D7 5

7 Link Budget Analysis: Antennae Design Strategy The Deep Space Network (DSN) will be used for earth-based reception of signals from the spacecraft and from the L2 relay satellite The DSN employs three parabolic antennae, each 34 meters in diameter, located such that either the spacecraft or satellite can communicate with one of the three at all times D7 6

8 Link Budget Analysis: Antennae Design Strategy The spacecraft must feature an omnidirectional antenna for UHF communication with astronauts during EVA because the astronauts may scatter about the vicinity of the spacecraft, and sustained reception of this signal by all astronauts is critical First design strategy: determine whether or not this omnidirectional antenna on the spacecraft can satisfy the transmission requirements of the spacecraft-to-earth and spacecraft-to-satellite links D7 7

9 Ku Band Spacecraft to Earth with Omnidirectional Transmitter Ku band frequency = 12 GHz Slant range = 4*10 5 km o Conservative estimate based on lunar distance at apogee Receiver system noise temperature = 300 K o Surface-based signal reception Diameter of receiving antenna = 34 m Receiver efficiency = 0.55 Diameter of transmitting antenna = λ/π = 8.0*10-3 m Transmitter efficiency = D7 8

10 Ku Band Spacecraft to Earth with Omnidirectional Transmitter D7 9

11 Revised Antennae Design Strategy Even at the maximum Ku band transmitter power of 20 W, the link margin is only db, which is far lower than the 3 db target Revised designed strategy: replace the omnidirectional transmitter antenna on the spacecraft with a small parabolic antenna (10 cm diameter) and determine the effect on link margin o Diameter of transmitting antenna = 0.10 m o Transmitter efficiency = 0.55 D7 10

12 Ku Band Spacecraft to Earth with Parabolic Transmitter D7 11

13 Ku Band Spacecraft to Earth with Parabolic Transmitter With a parabolic antenna on the spacecraft, the desired 3 db link margin can be obtained by the Ku band spacecraft to earth link with less than 0.3 W of transmitter power At the maximum transmitter power of 20 W, the link margin is an astounding db D7 12

14 S Band Spacecraft to Earth S band frequency = 2.5 GHz Slant range = 4*10 5 km Receiver system noise temperature = 300 K o Surface-based signal reception Diameter of receiving antenna = 34 m Receiver efficiency = 0.55 Since a parabolic antenna must be present on the spacecraft for Ku band transmissions, we may as well use the same antenna for all non-eva transmissions o Diameter of transmitting antenna = 0.10 m o Transmitter efficiency = 0.55 D7 13

15 S Band Spacecraft to Earth Transmitter power required to meet target link margin = 6.4 W D7 14

16 Ka Band Spacecraft to L2 Relay Satellite Ka band frequency = 32 GHz Slant range = 6*10 4 km o Distance from moon to earth-moon L2 Point Receiver system noise temperature = 46.4 K o Deep space signal reception Diameter of receiving antenna = 0.5 m o Conservative approximation Receiver efficiency = 0.55 Diameter of transmitting antenna = 0.10 m Transmitter efficiency = 0.55 D7 15

17 Ka Band Spacecraft to L2 Relay Satellite Transmitter power required to meet target link margin = 0.7 W D7 16

18 Ku Band L2 Relay Satellite to Earth Ku band frequency = 12 GHz Slant range = 4.6*10 5 km Receiver system noise temperature = 300 K o Surface-based signal reception Diameter of receiving antenna = 34 m Receiver efficiency = 0.55 Diameter of transmitting antenna = 0.5 m Transmitter efficiency = 0.55 D7 17

19 Ku Band L2 Relay Satellite to Earth Transmitter power required to meet target link margin = W D7 18

20 UHF Spacecraft to Astronauts during EVA UHF frequency = 0.9 GHz Slant range = 20 km o Conservative estimate, given that the maximum distance the Apollo Lunar Rover traveled from the Lunar Module was 5 km Receiver system noise temperature = 300 K o Surface-based signal reception Diameter of receiving antenna = λ/π = m Receiver efficiency = Diameter of transmitting antenna = λ/π = m Transmitter efficiency = D7 19

21 UHF Spacecraft to Astronauts during EVA Transmitter power required to meet target link margin = W D7 20

22 Power Requirements Below are the transmitter powers required for each link to achieve a 3 db link margin Link Ku Band Spacecraft to Earth 0.3 S Band Spacecraft to Earth 6.4 Ka Band Spacecraft to L2 Relay Satellite 0.7 Ku Band L2 Relay Satellite to Earth UHF Spacecraft to Astronauts during EVA Transmitter Power Required (W) D7 21

23 Antenna Design The 10 cm in diameter parabolic antenna will be placed in the unpressurized capsule tip Its mount provides rotational degrees of freedom in order to point the antenna in any direction necessary D7 22

24 Sensor Top-Level Block Diagram D7 23

25 Sensors List Sensor Type Sensor Quantity Criticality Atmospheric Temperature Humidity Carbon dioxide/monoxide Oxygen Pressure (with time) Life Critical Non-Critical Life Critical Life Critical Life Critical Sampling Frequency 10 Hz 5 Hz 20 Hz 50 Hz 40 Hz Dynamics and Control Accelerometer Gyro Laser Rangefinder Mission Critical Mission Critical Mission Critical 3 khz 8 khz 1 khz Miscellaneous Camera Clock (Computer Core) Radio Receiver Voltage Current Non-Critical Mission Critical Mission Critical Mission Critical Mission Critical 60 Hz 200 MHz 100 MHz 1 MHz 1 MHz D7 24

26 Criticality Definitions Life Critical: loss or malfunction of this sensor could result in loss of life of crew member(s) Mission Critical: loss or malfunction of this sensor could result in failure of mission goals Non-critical: not necessary for either life support or mission success but may contribute to astronaut comfort D7 25

27 Sensor-Based Systems Atmosphere o Carbon dioxide sequestration system o Oxygen valve o Radiator controls Dynamics and Control o RCS thrusters o Main descent/ascent engine Miscellaneous o Speakers/headphones o Computer screens o Power distribution control D7 26

28 Design/Build/Test/Evaluate (DBTE) Concepts Our DBTE concepts are intended to test several human factors issues with the capsule design that cannot be easily predicted These tests will examine the ease of astronaut operations in the capsule as well as critical mission maneuvers that require astronaut control The following concepts are revealed in order of decreasing priority D7 27

29 Concept 1: Suiting Up Research objective: o Is there sufficient room inside the cabin for the crew to put on spacesuits Required mockup/test apparatus: o Simple mockup of cabin interior, with emphasis on accurately representing capsule volume Material can be cheap (e.g. cardboard, fabric, etc.) as long as dimensions are correct D7 28

30 Concept 1: Suiting Up Concept of test operations: o Have 3 people enter cabin mockup with full sized suits and put them on o Image shows the process of suiting up (one crew member in suit, one member aiding another) D7 29

31 Concept 2: Ingress/Egress Research objective: o Can crew members efficiently enter and exit the crew cabin in microgravity and partial gravity situations o Is current ingress/egress design feasible Required mockup/test apparatus: o 2 m x 1 m aluminum door frame with operational hatch (including tension rods) o Deployable ladder with door extends down to the surface (2.5 m) o Ballasted underwater simulation at 0 g and 1/6 g in spacesuits D7 30

32 Concept 2: Ingress/Egress Concept of test operations: o Verify that crew can fully open/close the hatch, deploy/retract the ladder and make it down to the surface/up from the surface within a reasonable time o Subject is a 95-percentile male fully suited Largest part of suit: 0.8m x 0.6m x 0.2m Portable Life Support System D7 31

33 Research objective: o Can crew maintain necessary standing posture to operate controls and use windows during lunar descent and landing Required mockup/test apparatus: o Simple workstation mockup (correct dimensions are focus) o o Concept 3: Lunar Descent Operations Need workstation desk area, control placement, window Cheap, waterproof materials - aluminum, wood Can use workstation mockup for Suiting Up concept to provide cabin with appropriate obstacles Ballasted weight mimicking maximum landing loads Time how long subjects can retain necessary posture and compare to landing time Find maximum acceleration under which subject can retain posture for time of landing D7 32

34 Concept 3: Lunar Descent Operations Concept of test operations: o Test how difficult it is to maintain posture (or how long posture can be maintained) and operate controls under extra weight and lateral forces o o Verify that crew can maintain maximum line of sight Ballasted weight of under optimum and emergency landing conditions acceleration D7 33 F = ma

35 Concept 4: Landing Site Visibility Research objective: o Determine landing site visibility depending on horizontal and vertical velocities during landing Required mockup/test apparatus: o Computer simulation Concept of test operations: o Test a range of landing profiles to determine landing site visibility area and time. Once mission planning has determined the landing profile, this will simulate the visibility of the landing site D7 34

36 References "Suit Characteristics." MX-2. University of Maryland Space Systems Laboratory. Retreived 8 December < roject/characteristics.html> D7 35