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MECH 372 / ENGR 372
Space Systems Design and Engineering II
 
Assignment #7 - Payloads Solutions
 
 
Objective: The objective of this assignment is to demonstrate an initial understanding of different payload classes and how they drive the design of the satellite bus and the overall space system and mission architecture. 

In this class, we've talked about different types of space missions, with four of the primary categories being Navigation spacecraft, Observation spacecraft (which may look at the Earth or out into space), Communication spacecraft, and Exploration spacecraft.

In this assignment, you will work on your own to learn more about these systems in order to supplement the presentations made during lecture.

 

NAVIGATION - During lecture, we will spend little time discussing the ways that navigation satellites work. These questions and their references will introduce you to satellite navigation technology. You should definitely work your way through the outstanding Trimble Navigation's GPS Tutorial; additional references include the Wikipedia SatNav Overview, the Global Positioning System Overview, and the Wikipedia Overview of GPS. Other references are certainly available via the web.

1. First, you should certainly be aware that satellite-based navigation is only one method of navigation; many others exist and have been used for far longer.  List three navigation methods that do not rely on spacecraft.

Landmarks, celestial, ded reckoning, LORAN radio navigation, etc.

2. As for satellite navigation services, you are certainly aware of GPS.  But there are others, and it is worth being aware of at least a little information about them:

i) What was the name of the first operational satellite navigation system, who used it, and over what time period did it operate? What type of orbit were the satellites in? What was the approximate accuracy of the system?   Briefly describe how a navigation solution was obtained.

TRANSIT, US Navy, 1960-1996. 5-10 satellites in polar LEO ~ 600 nm. 1 satellite had to be in view to get a position fix. Often, none were in view.100-200 meters accuracy when available.  To obtain a position fix, the Doppler s-curve was obtained by listening to the satellite, and the shape was analyzed to determine position on the Earth.

ii) What country runs GLONASS?  How many satellites are in the system, and in what orbits are they? How accurate is the system's Standard Precision (SP) service? 

Russia. ~24 in 3 MEO (12-hour period) orbit planes. ~70 meters.

iii) What country(ies) is (are) developing Galileo?  When will this system become operational? How accurate is the system intended to be? Why is this system being developed given the global availability of GPS?

European Union. Slated for operation in 2020. Accurate to a meter. Developed in order to have a positioning system independent of US in case of political disagreement.

3. The original operational GPS system used 24 satellites in 6 different orbital planes (although more satellites are now operational). 

i) Why are so many satellites used and deployed in different planes? 

In order to achieve complete global coverage with multiple satellites (5 or 6) in view of any location at one time.  

ii) GPS receivers use "trilateration" in order to determine their location in space. Ideally, 3 trilateration measurements would be required to fix a point in space. GPS, however, uses 4 such measurements. What is the primary reason for doing this?

4 measurements provide an accurate fix even with inexpensive clocks in the receivers (needed to make them practical in terms of cost).

iii) One technical demand of the GPS system is high precision maintenance of the orbits of the individual spacecraft (which also requires high precision sensing of their position using techniques other than GPS). Explain why this is critical by describing how this requirement leads to good performance in terms of a GPS receiver being able to determine its position.

Trilateration measurements are from the "known" locations of the satellites. Any errors in these locations lead to solution errors. So, high orbits (away from affects of drag) are used, satellites are maintained in precise orbits through precision tracking and use of propulsion, and any errors from their predicted locations are immediately distributed to the receivers via the GPS signals that are transmitted to the ground.

iv) What is GDOP and why does the geometric relationship between the receiver and the transmitters make a difference in the precision of the receiver's position estimate?

Geometric Dilution Of Precision

There are errors in using time to determine the actual distance from a satellite.  The manner in which the errors from multiple satellites add together depends on the different directions from which the signals arrive, and therefore the geometry of the receiver/satellite system.  If more than 4 satellites are in view, choose the satellites that give the best relative geometry... the lowest "dilution" of precision.

  

4. To quantitatively explore the concept of multi-lateration, let's do a few simple 2-D "bi-lateration" planar examples.  Imagine that there are two signal sources, each broadcasting a signal that moves at 1 unit per second.  One source is at the origin of the Cartesian plane, (0,0).  The second source is at (10,0).  

i) Your receiver states that you are 7.07 seconds away from both signal sources.  What are your possible locations in the Cartesian plane?

(5, 5) and (5, -5)

ii) Your receiver states that you are 4.24 sec from the first source and 7.62 sec from the second source.  What are your possible locations in the Cartesian plane?

(3,3) and (3,-3)

iii) Your receiver states that you are 13.42 sec from the first source and 12.65 sec from the second source.  What are your possible locations in the Cartesian plane?

(6,12) and (6,-12)

   

REMOTE SENSING - These questions relate to the phenomena, technologies and associated system sizing relationships for remote sensing missions. A primary reference for answering these questions is Section 11.2 of Understanding Space.  Other references exist, of course, and you are encouraged to seek them out as necessary.

4. In the context of remote sensing, briefly explain the difference between a passive and an active sensing system.

Active provides its own energy that is transmitted, reflected and received. A passive receives energy provided by a different source, perhaps visible light reflected by the target or infrared energy radiated by the target.

5. How do high-performance infrared sensors drive the thermal design of a satellite, and why is this the case?

Sensors must be cooled to 77-120 deg K in order to be sensitive enough to detect small amounts of infrared energy. This is cold, and it often requires cryogenic thermal design approaches, liquid nitrogen, active refrigeration, etc.

6. You have an option of putting your remote sensing satellite into either an LEO orbit (altitude of 600 km) or a geosynchronous orbit (altitude of about 35,800 km).  At geosynchronous altitude, your camera system is capable of 10m resolution.  What resolution would the same camera system achieve in the LEO orbit?  Refer to Eq 11-8 in Understanding Space.

As per Eq 11-8, the ratio of resolutions is the ratio of altitudes s.t. 450/35800 = 0.0168.  So, if the geo satellite had a resolution of 10m, the LEO satellite would have a resolution of 0.168m (16.8 cm).  You could also say that the LEO resolution was nearly 60 times better (1/.0168=59.67).

7.  Many hot objects emit infrared energy and are interesting enough that we want to find and track them from space; things like forest fires, rocket plumes, etc.  Given the information in the Sellers chapter as well as that from the thermal subsystem lectures about radiation, consider a object that is 1,000 deg C., and for the purposes of this problem assume that it is a perfect blackbody radiator.  MIND YOUR UNITS!!!

  1. What is its radiated power per m2 of external area? 

J=sigma*T4 = 149 kW/m2

  1. What is its wavelength of peak emission?

lambda = 2898/T microns = 2.28 microns

  1. If you were to design an observational payload tuned to this wavelength (from part (b)) with 5 meter resolution at an altitude of 450 km, what aperture diameter would be required?

D=2.44*lambda*alt/res = 2.44*.000,002,28*450,000/5=0.5 m

8. Briefly describe how the design of a remote sensing system effects other subsystems for the following situations:

  1. Consider an active microwave radar.  Compared to a visible camera with a 1m instrument resolution, how does the radar payload affect the overall mechanical design and configuration of the satellite?

Wavelength is ~500,000 times greater, which means the aperture diameter is ~500,000 times greater.  This is huge, leading to enormous radar apertures - so BIG systems - since it isn't feasible to scale that much, radar satellites typically operate at lower resolutions and/or use active synthetic aperture approaches to establish an effective diameter that is much larger than the physical size of the satellite's antenna (the effective diameter is associated with the distance the satellite travels during the flight time of the radar signal - in addition, images are often created using multiple 'snapshots')

  1. How would the communication subsystem be impacted by a mission that required large quantities of high resolution, realtime imagery?

Downloading this amount of information in realtime is an enormous driver, affecting the data rate of the communications link, the availability of ground stations, probably low Earth orbit which means many stations around the world to ensure realtime connectivity, etc.

  1. Would you expect a high precision remote sensing spacecraft to use a spin-stabilization attitude control technique? Why or why not?

Probably only with a despun platform. More likely, it would be a 3-axis stabilized satellite with precision pointing.

 

COMMUNICATION - The questions below relate to major design considerations when developing high power communication missions.  Relevant references are noted below.

9. High power communications spacecraft need to generate a LOT of power, and they thermally dissipate a LOT of power.  From the thermal perspective, consider a thermal balance for a spherical communication spacecraft with a 1 meter diameter in LEO. Following the white paint example shown on Slide 42 of the Thermal Subsystem lecture, the steady state temperature with 0W of internally dissipated power was -67 deg C. 

  1. Now, assume that the satellite dissipates 20 kW of power. Rerun the analysis and find the new steady state temperature.   Assume e=0.9.  Is this a significant change that would require a complete rethinking of the thermal design?

341 deg C = 614 deg K , yes, of course it is a significant change!

  1. With white paint, the absorptivity to emissivity ratio is already close to a practical limit.  Another strategy is to increase the area that the satellite radiates to space in order to lower the steady state temperature.  What would the spherical diameter need to be in order to achieve a satellite temperature of 300 deg K given the internal dissipation of 20 kW of power? Is this a significant change?   Think about what this implies for the size of satellites (and/or their radiators) that are very high power.

4.6 meters or so.  

This is a very significant impact on launch volume, moving heat to the external radiator surfaces, structural design, etc.

10. The Link Equation, shown below, relates the quality of a communication link (in terms of signal to noise ratio) to many of the driving contributions such as broadcast power, antenna gain, losses over distance and through the atmosphere, etc.

where

  • E/N is signal to noise (we want this to be large)
  • Pt is transmit power
  • Gt and Gr are the antenna gains of the transmit and receive antennae, respectively, and they are proportional to the area of the antenna for broadcasts at a given frequency
  • Ls is space loss, which is proportional to 1/(distance)2 for broadcasts at a given frequency
  • La is attenuation loss due to absorption in atmosphere, etc.
  • Ll is line loss through transmission wires
  • k=Boltzmann's constant
  • Ts is the system noise temperature
  • R is the data rate

More information on the Link Equation is found in the MECH 371 presentation supplement on this subject.  There are many things to learn regarding where this equation comes from and how it is used in the design and analysis of communication systems.  However, using this equation as a rough guide, we can determine some of the big trade-offs when it comes to designing a communications payload (as well as the communications subsystem for a satellite with a different payload).

  1. If the altitude of the satellite is cut in half, what is the effect on E/N when the satellite is directly overhead (assuming all of the parameters remain the same)? [increases by a factor of 4]
  2. If you need to double the E/N, what could you do with each of the following parameters, and if you did this, what would the "cost" be in terms of bus subsystems or the mission?
    • the transmit power - double the power, but this may require you to dramatically increase power generation and storage; may also increase radiator size since power dissipation goes up
    • the data rate - you could cut this by a factor of 2, but this may affect the quality of the mission and quantity of data products
    • the satellite antenna area (you will need to understand the relationship between antenna area and gain) - double the area, but this would require a larger antenna which drives structural design, mechanical configuration, mass for propulsion, inertia for attitude control, etc.

 

SCIENCE - These questions relate to a generic planetary spacecraft, which for our purposes represents a "science class" satellite mission.  

11. As with other spacecraft, the requirements and characteristics of planetary missions drive the design of the bus subsystems/components. For a generic planetary mission, consider how various subsystems of the planetary craft would differ significantly compared to those of a satellite designed for a low Earth orbiting mission.  

For example, for propulsion (which we study in MECH 371), planetary spacecraft often require significant fuel and more substantial thrust systems (compared to an LEO mission) for achieving its interplanetary trajectory - this increases mass, requires thermal management, etc.

For each of the subsystems listed below, provide a few similar thoughts on the primary effects and considerations:

  1. Power generation - significantly different distances from sun - so perhaps great solar generation at Mercury or Venus, or perhaps much worse thereby requiring RTGs other other generation mechanisms
  2. Attitude sensing and control - we often use properties of the Earth for sensing (IR signature of Earth, Earth's magnetic field, etc.) and control (magnetic field, etc.); this would have to be modified appropriately
  3. Communications systems, both on-board and on the ground - longer than normal distance requires large dishes (a la the DSN) on the ground
  4. Thermal - doesn't have the Earth albedo or IR input which often is used to help keep things warm using passive techniques during cruise phase to planet; direct solar would vary dramatically and depend on whether you're getting closer or farther away from the Sun; would need to know IR and albedo environment of the new planet once you get there
  5. Structures - Since you may need large antenna, RTGs, propulsion units, etc., there would be significantly different demands on the structure in terms of accommodating large assemblies, etc.