MINRAD Research on Miniaturized Radiometry

1
MINRADResearch on Miniaturized Radiometry

• Critical Design Review
• 13 January 2000

2
Agenda

• 0930 Welcome and Opening Remarks
• 0940 The MINRAD Program
• 1000 Radiometer Optics
• 1040 Radiometer Mechanical Design
• 1200 System Electronics
• 1240 Payload Control
• 1320 Schedule and Program Status
• 1330 Closing Review Action Items

• Dr. James Russell III / HU
• Dr. Scott Bailey / HU
• Dr. Mark Larsen / USU
• Dr. Ralph Haycock / USU
• Mr. Wayne Sanderson / USU
• Mr. Timothy Green / HU
• Ms. Gbemi Munis / HU
• Mr. Richard Wright / HU
• Mr. Peter Mace / USU
• Mr. James Ulwick / USU

Informal group discussions following the meeting,
1400-1500, or as time allows.
3
The MINRAD Program

• Dr. Scott Bailey
• Center for Atmospheric SciencesHampton
  UniversityHampton, VA 23668
• Phone [phone]Fax [phone]
• E-mail Scott.Bailey_at_HamptonU.EDU

4
Program Goals

• We wish to develop advanced, innovative,
  miniature, infrared sensors for investigations of
  atmospheric infrared backgrounds.
• We will actively involve students and faculty
  from Hampton University and SDL/USU in research
  and technology using spaced-based sensors.

5
Scientific Goals

• Provide improved understanding of the
  relationship between eddy diffusion and airglow
  emissions by performing sounding rocket
  measurements of O2(1?) emissions, OH Meinel
  emissions, and electron density.
• temporal changes, baseline OH chemistry,
  turbulent velocity, energy dissipation rate, and
  eddy diffusion processes
• Perform synergistic science studies of temporal
  and spatial airglow variations using the
  combination of MINRAD and TIMED SABER satellite
  data.

6
The Team
7
The Team

• As an affiliated research center for BMDO for
  research and technology in areas of sensors and
  space surveillance, Space Dynamics Laboratory
  (SDL) is leading the way in miniaturizing the
  radiometer for this project. PI Mr. James Ulwick
  is a pioneer in upper atmospheric exploration. He
  is leading the miniaturization initiative to help
  BMDO reach its goals and objectives through new
  technologies at lower costs.
• Hampton University (HU) has two leading
  atmospheric scientists committed to measuring and
  understanding the mechanisms exciting atmospheric
  species. PI Dr. James Russell III has extensive
  experience at NASA, has been PI of numerous space
  borne IR remote sensing instruments, and is
  currently PI of the SABER experiment on the TIMED
  spacecraft. HU, a historically black
  college/university (HBCU), is committed to
  developing a premier program in atmospheric
  sciences. Co-I Dr. Scott Bailey has successful
  experience leading students on NASA remote
  sensing missions.
• SPAWAR Systems Center San Diego - Mission is to
  be the pre-eminent provider of integrated C4ISR
  solutions for warrior information dominance. Dr.
  Clifton Phillips, assistant technology agent for
  multisensor tracking and optical target
  characterization, is monitoring this work.

8
BMDO Goals and Objectives

• The Ballistic Missile Defense Organization
  (BMDO) is tasked with providing a missile defense
  system to protect the U.S., its forces deployed
  abroad, and its friends and allies against
  accidental, unauthorized, and limited ballistic
  missile strikes. The agency has three mission
  focus areas.
• The first priority is to develop and deploy
  increasingly capable Theater Missile Defenses
  (TMD) to meet existing missile threat to deployed
  U.S. and allied forces. (Emerging threats include
  cruise missiles.)
• The second priority, as a hedge against the
  emergence of long-range ballistic missile
  threats, is to develop options to deploy a
  National Missile Defense (NMD) for the United
  States.
• The third priority is to continue supporting
  research on more advanced ballistic missile
  defense technologies to keep pace with the threat
  and improve the performance of theater and NMD
  systems.

9
Technical Objectives

• Develop radiometers that take advantage of orders
  of magnitude lower cost delivery systems.
• This makes atmospheric research routine and
  cheap.
• Small size and weight goals make the miniaturized
  radiometer sensor attractive for sensor suites
  aboard other platforms such as aircraft, UAV,
  extended seeker capabilities, and portable ground
  stations.
• Exploit development campaign for low cost Viper
  IIIA/DART delivery system by using it as
  development platform and baseline packaging
  constraint.
• Viper/DART serves as the role model in reduce
  sounding rocket lifecycle cost.
• Development of the experiment launch platform is
  funded by SDL internal research and development.
• Package two-channel radiometer into a two-inch
  diameter payload compartment.
• Develop miniaturized sensor management and
  housekeeping functions.
• Facilitate easy growth path to cryogenic upgrades.

10
Science Objectives

• Contribute to the overall atmospheric optical
  characterization goals of BMDO.
• Quantitatively test temporal changes in the
  relationship between O2(1?) airglowand O3
  concentration.
• Address the baseline chemistry of OH using rocket
  and ground based measurements.
• Derive turbulent velocity, energy dissipation
  rate, and eddy diffusion coefficients in middle
  atmosphere using spectral information from
  electron density fluctuations.
• Study relationship between eddy diffusion and
  airglow using photometer, radiometer and atomic
  oxygen measurements compared to 1-D kinetic
  transport model predictions.
• Perform synergistic science studies and define
  temporal changes using miniature radiometer
  measurements and similar measurements from
  complementary NASA spacecraft (i.e. TIMED).
• Extract clear understanding of atmospheric
  generated clutter.

11
How MINRAD Helps BMDO

• Directly addresses BMDOs need to identify and
  understand atmospheric IR background clutter.
• Develops miniaturized sensor technology for the
  measurement of clutter.
• Enhances scientific understanding of the
  processes creating clutter.
• Miniaturized and rugged radiometry expands the
  range of available launch vehicles allowing
  rapid, inexpensive, and reliable access to space
  for the testing of new sensor technology.
• Supports upgrade path for interceptor, seeker,
  and space based sensor technologies (miniaturized
  radiometry provides reduced weight and presents
  minimal impact on maneuverability and
  performance).
• Potential use for verification of background
  clutter in local theater of operations.

12
Development Path(Smaller-Lighter-Cheaper)
BMDO Capability
Surveillance Sensor
Miniature Cryogenic Sensor (l,mass optimized)
Technology / Hardware Trend
Miniature Cryogenic Sensor (LWIR)
Miniature Near IR Sensor (current work)
Large Cryogenic Sensor (large mass)
Time
Current Technology
MINRAD Phase III
Broadband Spectral Ability
Enhancement Infusion
MINRAD Near Term
13
MINRAD Exploits Developed Low cost Launch Vehicles

• Provides significant cost reductions for science
  platforms because smaller and lighter payloads
  can be flown on lower cost launch vehicles.
• Lower cost science platforms allow more
  experiments to be flown to understand
  temporal/spatial processes that create emissions
  and provides expanded capability to meet
  multi-spectral needs.
• Miniaturization technologies can be implemented
  now with low cost and risk.

10000
1000
100
Launch Vehicle Cost Relative to Viper DART
10
1
Terrier Malamute
Black Brant
Viper DART
Delta II
Pegasus
14
Sample Flight Profile
15
Radiometer Optics

• Dr. Mark Larsen
• SPACE DYNAMICS LABORATORYUtah State
  University1695 North Research Park WayNorth
  Logan UT 84341-1947
• Phone [phone]Fax [phone]
• E-mail Mark.Larsen_at_SDL.USU.EDU

16
Outline

• Optical Design
• Model
• Optical, radiometric, electrical
• Sensitivity
• Detectors, NEP, SNR
• Background
• Instrument, solar, nose cone

17
Optical Design Objectives

• Package length lt 5 in.
• Package footprint lt 1.5 in. x 1 in.
• Weight lt 1 lb.
• Acceleration load of 150 gs
• Unknown thermal shock
• SNR 50
• Full field of view 15 (achieve SNR)

18
Mechanically ChoppedDouble Barrel Radiometer

• Features
• Two optical barrels
• Simple filters
• Advantages
• Low technology risk
• Disadvantages
• Moderate packaging difficulties

19
Optical Design Characteristics

• Clear aperture 10 mm
• Full field of view 12
• Two spectral bands
• Band 1 1.263 ?m - 1.290 ?m (?? 27 nm)
• Band 2 1.463 ?m - 1.563 ?m (?? 100 nm)
• No vignetting
• Uniform detector irradiance
• Transmittance of optics 0.56
• Diameter of the detector 2 mm

20
Optical Design Challenges

• Packaging
• Lens mounting for 150 g
• Baffling

21
Optical Layout
22
Band 1 Spot Diagrams
23
Encircled Energy
24
Radiometer Overview
25
Radiometer Baffle
26
Radiometric Model
27
Electrical System

• Design to respond to frequencies higher than the
  spin rate of the rocket (20 Hz)
• Electrical noise bandwidth 50 Hz
• Chopping frequency 300 Hz
• EGG InGaAs thermoelectrically cooled detectors
  operating at 10 ºC
• Size 2mm, cutoff 1.7 mm, typical D 2.1e13
• NER 3.8e-12 w/cm2sr
• 362 Rayleighs in band 1
• 305 Rayleighs in band 2

28
Instrument Background

• For a temperature of 20 ºC inside the sensor, the
  background is
• 6 in band 1
• 12 in band 2

29
Background From Nose Cone

• Assuming the nose cone has a temperature of 200
  ºC, the nose would cause an error of 5 at 50
  meters if it is in the field of view.
• Since we wont know the range to the nose, we
  need to get it out of the field of view.

Band 1, Minimum cross section
30
Non-Rejected Solar Radiance

• Given
• A coning angle of 5º, and a mean angle to the sun
  of 80º.
• The relative response of sensor at 80º is 10-8.
• Then the in-band non-rejected solar radiance
  varies between approximately 12 and 30 of the
  signal level in both bands.

31
Summary

• SNR of 50 using existing detector technology.
• Background Signal (Noise) Reduction
• Get the nose cone out of field of view.
• Keep telescope optics and baffling at 20 ºC or
  less.
• Make angle to sun as large as possible and coning
  as small as possible.
• The optical design satisfies the requirements.

32
Radiometer Mechanical Design

• Dr. Ralph Haycock
• SPACE DYNAMICS LABORATORYUtah State
  UniversityUMC 4130Logan UT 84322-4130
• Phone [phone]Fax [phone]
• E-mail Ralph.Haycock_at_SDL.USU.EDU

33
Topics

• Mechanical Assembly
• Payload Mounting
• Component Identification
• Component Mounting
• Isometric Views
• Exploded Views
• Section Views

• Mechanical Analysis
• Chopper blade
• Stresses due to
• Acceleration or Thrust
• Spin
• Vibration
• Thermal Analysis
• Future Testing and Analysis

34
Payload Mounting
Radiometer
35
Component Identification
36
Component Mounting
37
Isometric (Fore)
38
Isometric (Aft)
39
Isometric Exploded (Fore)
40
Isometric Exploded (Aft)
Motor/Shaft Coupling
41
Section Views
Section A-A Section
B-B Section C-C
42
Analysis of theRadiometer Chopper Blade

• Axial acceleration or g loading
• Radial acceleration or spin loading
• Vibration
• Resonant frequencies
• Mode shapes

43
Prototype Disk Analysis

• Material
• Aluminum 6061-T6 was used for the analysis and
  testing.
• Other material (stainless steel, spring steel,
  etc.) may be considered in the final design.
• Acceleration g loads
• Validation by measuring displacement under a test
  load.
• Rotating Loads
• Validated with a closed form solution of a solid
  disk.
• Vibration Loading
• Validation by identifying the resonant
  frequencies in a prototype model.

44
Acceleration or g Loading

• Contact between the chopper disk and structure
  may result in no chopper rotation
• Material
• 6061-T6 Aluminum
• Mass Density 26.6 KN/m3 (0.098 lb/in3)
• Acceleration load of 150 gs during launch

45
Finite Element (FE) Model Displacement

• Found in the z direction (normal to the surface
  of the disk)
• Coarse and fine mesh FE models created to verify
  convergence
• Will be verified with actual measurements

46
DisplacementCoarse Mesh FE Model
47
DisplacementFine Mesh FE Model
48
Finite Element (FE) ModelStress

• The stress level shown is for the top of the thin
  shell element
• Must be less than the fatigue stress for the
  dynamic load and less than the yield stress for
  the g loading
• For 6061-T6 Aluminum
• Yield strength Sy 296 MPA (43 kpsi)
• Fatigue stress St 90 MPA (13 kpsi)

49
StressCoarse Mesh FE Model
50
StressFine Mesh FE Model
51
Results

• Displacement
• Maximum displacement occurs at the outer edges
  and is approximately 0.000283 inches
• Stress
• Maximum occurs near the disk center
• Maximum stress of 1.1 kpsi is well below the
  limit of 43 kpsi

52
Spinning Blade Analysis

• A disk rotating at 9000 RPM or 150 Hz induces
  body forces within the disk.
• Displacement could lead to interference with
  radiometer housing.
• FEA was used to determine stress and displacement
  for a thin disk with a hole in the center.
• The disk was meshed as quadralinear Mindlin shell
  elements using the IDEAS software.
• Partial validation has been obtained through a
  similar FEA with a closed form numerical solution.

53
FEA of a Spinning Disk
54
Closed Loop Solution for a Spinning Disk
55
Analysis Results for a Spinning Disk

• The stress level for the exact solution and the
  FE model are very close.
• Both solutions used the same boundary conditions.
  
• Mesh refinement will reveal a closer relationship
  between these two solutions.
• Coarse mesh of the entire disk was compared to a
  half section fine mesh model.
• The entire disk model and half model where
  compared to give validity to the fine meshed half
  sections.

56
DisplacementWhole Disk Coarse Mesh FE Model
57
DisplacementHalf Disk - Fine Mesh FE Model
58
StressWhole Disk Coarse Mesh FE Model
59
StressHalf Disk - Fine Mesh FE Model
60
Results

• Displacement
• Max radial displacement of 9.11E-6 inches
• Stress
• Maximum von Mises stress of 241 psi
• Well below maximum yield stress of 43 kpsi and
  the fatigue strength of 13 kpsi

61
Vibration Analysis

• Examine the resonant frequencies in the axial or
  z direction of the chopper disk.
• FEA for initial predictions
• Test set up
• Refine the FE model

62
Initial FEA of Disk

• Three fundamental mode shapes predicted
• First resonant frequency at 2040 Hz
• Second resonant frequency at 2600 Hz
• Third resonant frequency at 2830 Hz
• NOTE In the following pictures showing the mode
  shapes,the FE model included the mounting shaft
  below and thesensing accelerometer on top of the
  disk.

63
First Fundamental Mode
64
Second Fundamental Mode
65
Third Fundamental Mode
66
FEA OutputFirst Resonant Frequency
67
Vibration Testing

• Verify the FE model and the analysis.
• A large, 8 diameter, scale model was developed
  to illustrate the vibration mode shapes.
• Accelerometers were placed on the outer edges of
  the disk.

68
Test Chopper Disk
Accelerometer 1
Accelerometer 2
69
Video Images From Vibration Tests
70
Vibration Test DataAccelerometer 1
71
Vibration Test DataAccelerometer 2
72
Interpretation of results

• A very prominent resonant frequency at
  approximately 200 Hz
• Less prominent resonance around 160 Hz
• An FEA of the scale model predicted resonant
  frequencies of 142, 155, and 193 Hz
• The FEA of the scale model and the data from the
  vibration test were in close agreement

73
Future Testing and AnalysisChopper Blade

• Locate the resonant frequencies in the mounting
  shaft
• Determine the required preload in the jewel
  bearing
• Determine the friction in the jewel bearing

74
Future Testing and AnalysisChopper Motor

• Torque curves
• Analysis of mounting cradle
• Vibration of motor and chopper assembly
• In the axial direction
• In the transverse direction
• Determine the startup time
• Moments of inertia
• Total friction of motor and jewel bearing
• g load on motor bearings due the mass of the
  chopper, shaft, and rotor

75
Future Testing and AnalysisRadiometer System

• Vibration Analysis and Testing
• Axial and transverse axes
• Determine the mounting and structural integrity
  of the optical system
• Find resonant frequencies
• g loading in axial and transverse axis
• Simulate thrust loading in a centrifuge

76
Thermal Analysis

• Numerical Modeling
• FE models in IDEAS
• Finite difference modeling using SINDA
• Issues
• Heat transfer due to aero heating on nose cone
• Internal heating due to
• Chopper (motor inefficiency, bearing friction,
  etc.)
• Detectors (TE cooling, electronics, etc.)

77
Description of FE ModelPreliminary Thermal Model
of Nose Cone

• 3D solid model of the nose tip and the nose cone
  body
• 10 section of model nose cone meshed
• Thin wall sections 0.006 square by 0.063 thick
• Initial conditions and boundary conditions
• Initial temperature of model set at 22 C
• Outside surface set at 500 C

78
Sample FE ModelPreliminary Model of Nose Cone
79
SINDA Analysis

• SINDA is a comprehensive finite-difference,
  lumped parameter (circuit or network analogy)
  tool for analyzing complex thermal systems.
• SINDA offers steady-state and transient
  solutions.
• SINDA allows a second approach to the numerical
  modeling for verification of FE model developed
  under IDEAS.

80
Future Testing and AnalysisThermal

• FE and SINDA Modeling
• Nose cone
• Examine and minimize heat transfer to radiometer
  housing, prior to nose cone ejection
• Motor
• Verify sufficient heat sink for motor operation
• Verify isolation to prevent motor heat from
  reaching detectors
• Detectors
• Model and optimize heat transfer to aid in
  detector cooling

81
Summary

• Tasks have been identified and team members have
  been assigned to program
• Structural analysis and prototype fabrication has
  started
• Thermal issues have been identified and the
  analysis has started
• Test hardware has been ordered and testing is
  currently underway on motors and bearings

82
System Electronics

• Wayne Sanderson
• SPACE DYNAMICS LABORATORYUtah State
  University1695 North Research Park WayNorth
  Logan UT 84341-1947
• Phone [phone]Fax [phone]
• E-mail Wayne.Sanderson_at_SDL.USU.EDU

83
Overview

• Payload Overview
• Electronics by Sections
• Tail section
• Aft section
• Mid/Battery section
• Forward section
• Instrument section
• Radiometer Electronics
• Summary of Instruments

84
Payload Layout
85
Design Goals

• Smaller
• Surface mount components
• Blind Assembly
• Modular design
• Plug and play
• Bulkhead and motherboard
• Keep Pin Count Down
• Common control lines
• Split controllers

• Low Power (Primary Cells)
• Select low power devices where possible
• Use DC/DC converter to improve efficiency
• Flexibility
• Modular PC boards
• FPGA controllers

86
Aft Section Layout
87
Transmitter Section

• Transmitter
• L3 Communications T-401S
• 1 watt output
• 2279.5 MHz (S-band)
• 800 Kbps
• Antenna
• Slotted tail fin coupled with RF matching network
  provides circular polarization
• Temperature Sensors
• Transmitter 0-150 ºC, 195 samples/sec
• Tail Skin 0-300 ºC, 195 samples/sec

88
Accelerometer/Pressure/Umbilical

• Accelerometer Z (Thrust)
• 150/-50 g
• 1562 samples/sec
• Accelerometer X1 (Off-axis)
• 100 g
• 1562 samples/sec
• Accelerometer X2 (On-axis)
• 25 g
• 1562 samples/sec

• Pressure
• 0-30 psi
• 195 samples/sec
• Umbilical Interface

89
Power PCB

• Internal/External Power Switching
• GPS Power and Interface
• System DC/DC Converter
• DC/DC Temperature
• 0-150 ºC
• 195 samples/sec

• Battery Bus Voltage
• Umbilical (analog)
• Real time safety monitor
• Battery Bus Voltage
• 23-33V
• 195 samples/sec
• Battery Bus Current
• 0-1A
• 195 samples/sec

90
GPS PCB

• Global Positioning Receiver
• Rockwell Jupiter model
• Data rate 1200 words/sec
• Pseudo range operation
• Reset position at launch for faster acquisition

91
Aft Instrument PCB

• Main Control
• PCM data formatter
• Payload controls
• FPGA based
• Pyrotechnic Release
• Firing circuit
• NASA design
• DC/DC Monitors
• 4.5 to 5.5V _at_ 195 SPS
• 11.5 to 12.5V _at_ 195 SPS
• -12.5 to -11.5V _at_ 195 SPS

• Temperature Monitors
• Printed circuit board
• Aft battery pack
• Bulkhead
• Signal Conditioning
• Temperature monitors
• Voltage monitors
• 3 axis magnetometer

92
Mid Section Layout
93
Mid Section Components

• Aft Battery Pack
• 28 V _at_ 1 AH
• Fore Battery pack
• 6 V _at_ 5 AH
• Magnetometer
• 3 axis _at_ 2 Gauss
• 195 SPS

• GPS Antenna
• 2 patch antennae
• Diametrically opposed
• GPS Preamplifier
• Temperature Monitor
• 0-150 ºC _at_ 195 SPS

94
Fore Section Layout
95
Pyrotechnic Mount PCB

• Forward Battery
• Internal power switch
• Voltage monitor
• Current monitor
• Signal Conditioning
• Temperature monitors
• Forward battery monitors

• Radiometer Monitors
• TE temperature
• TE voltage
• Motor voltage
• Motor current
• Motor speed
• Mount for Release Pyro

96
Forward Instrument PCB

• Forward Multiplexer
• Forward A/D Converter
• PCB Temperature
• 0-150 ºC _at_ 195 SPS
• Forward Skin Temperature
• 0-300 ºC _at_ 195 SPS

• Radiometer
• Signal conditioning
• Motor control
• TE control
• DC Probe
• Secondary Digital Controller

97
Instrument Section
98
Instrument Section

• DC Probe
• Ring antenna
• Preamplifier
• Radiometer Detectors
• InGaAs detectors
• Preamplifiers
• TE cooler

• Radiometer Instrumentation Amp
• Chopper Motor/Encoder
• Temperature Monitors
• 0-150 ºC _at_ 195 SPS
• Baffle
• Motor
• TE Heat sink

99
Radiometer Electronics
100
TE Temperature Control
101
Chopper Control/Rectifier Sync
102
Digital Controls

• Launch Recognition
• Power Switching
• Digital Monitors
• A/D Control
• SYNC/SFID
• Chopper Control
• Rectifier Signal Delay

• Serial Data Output
• Bi-?-L Control
• GPS Control
• UART initialization
• DATA control
• PCM Control
• Data matrix
• MUX control
• A/D enable

103
Payload/Instrument Summary
104
Projections

• Through the use of surface mount devices and
  programmable logic, board space required for DART
  electronics is approximately 36 in2.
• Runtime for the battery packs (internal
  power).
  (Based on 40 ºC detector temperature)
• Digital controls are easily reconfigured through
  the use of programmable logic, limiting the need
  to redesign the PC boards.

105
Payload Control

• Hampton UniversityCenter for Atmospheric
  SciencesElectrical Engineering Department
• Timothy Green
• E-mail temitope39_at_hotmail.com
• Charles Hill
• E-mail Charles.Hill_at_HamptonU.EDU
• Gbemi Munis
• E-mail gbemisola_at_hotmail.com
• Richard Wright
• E-mail nrplongisland_at_hotmail.com

106
Objectives

• Hampton University will design and implement the
  software for embedded control of the rocket
  payload. This includes data sampling and
  formatting, chopper motor control, and handling
  of the payload telemetry.
• The software will be programmed into a Field
  Programmable Gate Array (FPGA) logic device.
• Hampton University will assemble a complete
  rocket payload assuming funds are made available.

107
Tasks

• Software Component Design Implementation
• Fore Section Embedded Logic Control
• Aft Section Embedded Logic Control
• PC-Based Decoder

108
Payload/Instrument Summary
109
Requirements
Digital Signals
Housekeeping
Science Sensors
Temperature Acceleration Pressure Battery Voltage
and Current Motor Voltage and Current
GPS, Digital Monitors
Optical Devices Langmuir Probe Radiometer
Field Programmable Gate Arrays
Chopper Motor Control
Telemetry
110
Telemetry Matrix
Bit Rate 800 Kbps Word Rate
50000 wps Sub-frame Rate 1562.500
sfps Major Frame Rate 195.310 fps
111
Field Programmable Gate Arrays

• Appropriate When Space Is Limited
• A large collection of discrete logical components
  are replaced by a significantly smaller package.
• Reduced Power Consumption
• FPGAs offer reduced power consumption as
  compared to boards populated by an equivalent,
  discrete logic circuit.
• Speed

112
Field Programmable Gate Arrays

• Reprogram in the Field
• Software enhancements may be added quickly and
  easily without removing the device from the
  rocket.
• Wide Product Selection
• There is currently a wide and flexible range of
  FPGA products commercially available.

113
Device Selection

• Altera MAX 7000 Device Family
• Free educational kit that includes a design
  laboratory package.
• Minimal resourcesneeded to get started.

114
Design Laboratory Education Board
115
Altera Hardware Description Language

• AHDL Uses Concurrent Logic
• AHDL is a text entry language for describing
  logic designs.
• AHDL is a concurrent language. Each line of
  source code is evaluated at the same time rather
  then sequentially.
• Examples of Source Code to Follow
• Logic Circuitry (e.g. AND Gate, OR Gate, etc.)
• Graphic Version of Source Code Is Always Possible

116
AND Declaration(AHDL Concurrent Logic)
SUBDESIGN GATE( a,b INPUT c OUTPUT)BEGIN
c a AND bEND
GATE
117
OR Declaration(AHDL Concurrent Logic)
SUBDESIGN GATE2( a,b INPUT c OUTPUT)BEGIN
c a OR bEND
118
2 to 1 Multiplexer(AHDL Concurrent Logic)
SUBDESIGN MUX2TO1( pressure, temperature,
address INPUT converter OUTPUT)BEGIN IF
address 1 THEN converter temperature ELSIF
address 0 THEN converter pressure ENDIFE
ND
119
Fore Section Altera
ALTERA
A/D
MUX 41
MUX 161
MUX 161
120
Aft Section Altera
Data From Fore Section Altera
MUX 41
A/D
Aft Altera
MUX 161
Serial Data Stream
SENSORS
GPS Data
121
Typical System Flow
Chopper Motor Driver and Encoder
Science
Optical Devices, Langmuir Probe, Signal
Output, Radiometer
Fore Altera
Sensor Sampling Sequence Control
Housekeeping
Temperature, Acceleration, Pressure
Analog MUX
A/D
Aft Altera
GPS
Read/Write
Transmitter
122
PC-Based Decoder

• Used for verification of data stream
• Recover of analog values in real-time
• Software/Hardware yet to be purchased (COTS)

123
Software Simulation

• Altera software provides I/O simulation
  capability
• software verification
• timing analysis

124
Chopper Motor Control
Chopper
Aft Altera
Fore Altera
Motor Encoder
Chopper Motor
Chopper Driver
Pulse Width Modulation
DC Voltage
125
Telemetry PCM Code

• Manchester Non-Return-to-Zero (NRZ).
• Also known as Bi-?-L (bi-phase L).
• Relatively inexpensive circuitry may be used
  throughout the system.
• It allows us to merge multiple signals on one
  common transmission line. (Time-Division
  Multiplexing).
• Other Options and Drawbacks.
• Unipolar NRZ requires DC couples channels.
• Polar NRZ requires negative and positive voltage
  supply.

126
Schedule and Program Status

• Peter B. Mace
• SPACE DYNAMICS LABORATORYUtah State
  UniversityUMC 4145Logan UT 84322-4145
• Phone [phone]Fax [phone]
• E-mail petemace_at_cc.usu.edu

127
Program Schedule
128
Radiometer Development
129
Fabrication of Payload
130
Software Development
131
Assembly and Calibration
132
Launch Schedule
133
Purchasing (COTS)
134
Program Status 1

• Mechanical
• Prototype layout has been developed
• Analysis and testing of prototype elements are
  underway
• Initial hardware has been identified and ordered

• Electronics
• Tasks have been identified
• Team members have been assigned tasks
• Board layout can be started
• Software development is underway

135
Program Status 2

• Optics
• Detectors have been identified and have been
  ordered
• Optical elements have been identified
• Lens and filter orders pending analysis of
  mechanical mounts

• Launch Schedule
• Proposed launch site
• NASA Wallops Flight FacilityWallops Island, VA
• Date dependant upon
• Number of payloads
• Science goals

136
Review and Action Items

• Mr. James UlwickPrincipal Investigator
• Steward Radiance Laboratory139 The Great
  RoadBedford MA 01730
• Phone [phone]Fax [phone]
• E-mail srl_at_tiac.net

• Dr. James Russell IIIPrincipal Investigator
• Center for Atmospheric SciencesHampton
  UniversityHampton, VA 23668
• Phone [phone]Fax [phone]
• E-mail James.Russell_at_HamptonU.EDU