No DX on LEO's ??? Create your own dream idea on a rainy night !
Have you ever looked back fondly at the DX you worked on AO40, AO13, or any High Earth Orbiting satellite, then looked in disappointment at the footprint of the typical 700km high Low Earth Orbiting satellite? Well it's Sunday here in London and having cut the grass walked the dogs and watched the rain start, I've decided to see if there's a practical solution for getting a satellite into an orbit that can offer better DX.
LEO satellites are 'easy' to launch as the 680  800km high orbit is one frequently used for Earth observation satellites. If you have the money, just hitch a lift on a convenient launcher going your way.
HEO satellites used by AMSAT have traditionally used a launch vehicle going to Geostationary Transfer Orbit. This type of launch is typically highly elliptical. The main 'passengers' are destined for geostationary orbit above the equator but AMSAT have successfully used 'kick motors' of around 400N thrust to raise the perigee to a safe height and extend apogee to 35,000 to 60,000km. These satellites offer great DX but are very complex. The typical AMSAT HEO spacecraft must have relatively high power transmitters to overcome the huge path loss over 50,000km, have electronics that is able to withstand the charged particles in the Van Allen radiation belts and generally weigh 80kg or more.
So, How about a satellite that is in a Medium Earth Orbit?? Or as I've called it  MEOSAT ?
A satellite in medium earth orbit would have excellent DX potential compared to a LEO. But how far above the earth should the satellite be?
Well, there is an area of space that presents a
logical choice. The Van Allen radiation belts are separated into two layers. ** The
lower layer is comprised of high energy protons between 600 and 6000km. The
second belt is essentially electrons and that occupies altitudes above 12,000km. So our MEOSAT could avoid damaging radiation by orbiting
in the "safe zone" between. 7000 and 11000km. *
Footprint of a Medium Earth Orbiting satellite at an orbital height of 7400km
By entering the satellite details into the
Nova tracking program you can see the characteristics of the orbit.
The plot above shows the International Space Station at 350km and 'MEOSAT' at
7400km
For my QTH in London a typical overhead pass lasts not for the 12 or 15 minutes
of a LEO satellite, but a full 90 minutes.
The DX potential is equally impressive:
California can work into Western Europe.
Italy can work Western Australia (just)
Northern Finland can work Capetown South Africa.
Looks good doesn't it ! MEOSAT will orbit the Earth 5.4 times each day. (mean motion)
So, what equipment would be needed to operate a satellite in this orbit? Let's compare AO51 AO40 at 50,000km and our fantasy MEOSAT.
Satellite
Height
Elevation Slant
range 2m path
loss 70cm path
loss
AO51
750km
15deg
2000km
142dB
151.4dB
AO40
45000km xxx
50000km
169.7dB
179dB
MEOSAT
7400km
25deg
9800km
155.5dB
165dB
UPLINK POWER
For a mode U/V transponder, the uplink to would be on 70cms and the downlink on
2m. To reach MEOSAT the path loss is 13.6dB greater than AO51 but 14dB
less than the path loss out to AO40. Without doing a full analysis
including receiver / antenna / path loss / filter and other losses we can
estimate the power needed by comparison with an AO40 station.
AO40......50W + 19 ele Yagi. = 2kW ERP.
MEOSAT = 14dB + 6 dB for simple antennas = 2kW  8dB =
300 W ERP (for a very good signal)
So a typical station maybe a 30 Watt radio and a small 70cm yagi of around 9
elements for a good SSB signal. 5  10W should work for CW.
DOWNLINK
A typical LEO with a 100kHz linear transponder  runs 1 Watt to a simple
antenna (1W ERP) and can be received on a small handheld beam of 2 or 3
elements, perhaps +4dB with good strength.
We've seen that MEOSAT has 13dB more path loss on 2m and so would need to
run 20 Watts to a similar antenna to be received on the same small beam.
However, 20W is too high for this idea! Other choices would be to reduce
the transponder bandwidth, use a directional antenna with some gain on
MEOSAT or increase the gain on the groundstation receive antenna.
To ensure the 'practicality' of the idea, lets reduce the bandwidth by 50% and
save 3dB. Then double the output power to 2 Watts saving another 3dB.
So now instead of our 13dB deficit we now have 13  3 3
= 7dB.
If we can increase the size of the groundstation antenna to add 4dB so the
required gain is now +8dB, that will leave us 3dB short and that could simply be
tolerated as a reduction in signal to noise ratio.  Again, a proper
analysis is required, but from the above 2W from a 50kHz transponder would make
the distance.
SATELLITE DESIGN
Sending any satellite to a Medium Earth Orbit is not a trivial task.
Launches are available to LEO at 800km or to GTO at 23000km. There is also the
possibility of the shuttle at 350km.
From this it can be seen that some form of propulsion will be needed.
The shuttle will not allow 'pyrotechnics' on board.
GTO is unsuitable at 23,000 and being 'separated' on the way up to GTO is
probably unrealistic. GTO launches are very expensive
So the only practical option is a LEO launch with propulsion and as with any
satellite the less mass you have, the less fuel you need to move it. Could
this be an opportunity for a small satellite with a low mass?
For several years now, Universities have launched
cubesats with varying degrees of success. The cubesat program produces
satellites for LEO launch having the general specification of 10 x 10 x 10cm and
having a mass of 1kg.
A recent project by students at the University of Delft has produced a satellite
design DelfiC3 which departs from the normal cubesat concept. The Delft satellite is apx
10 x 10 x 30cm. three times the length / volume of a normal cubesat and with deployable thin
film solar panels.
If MEOSAT were constructed as essentially a single function satellite it would
be entirely possible using modern SMD technology to build the transponder and
other electronics into a
10 x 10 x 10cm or other small space. If the Delft model http://www.delfic3.nl/
were followed, the extra volume would
allow enough solar cells to power the transponder. The extra volume provided in
a 10 x 10 x 30 structure could hold a small chemical or gas based propulsion
system. With 50% of the satellites volume devoted to propellant
and only a small mass to propel, this combination may well be enough to get
MEOSAT into a usable higher orbit..
The Van Allen radiation
belts and typical satellite orbits. Key: GEO—geosynchronous orbit; HEO—highly
elliptical orbit; MEO—medium Earth orbit; LEO—low Earth orbit. (Illustration
by B. Jones, P. Fuqua, J. Barrie, The Aerospace Corporation.)
http://www.aero.org/publications/crosslink/summer2003/02.html
Q1)  Anyone out there care to run through the necessary thrust equations to prove the idea?? final mass 1.5kg initial mass ?? 4kg ??
* Ref: Effects of the CME of oct 2003 on the safe zone. http://www.nasa.gov/vision/universe/solarsystem/safe_zone.html
** Ref: Nature / Physics Sept 2005. Tightening the radiation belts by Paul Hanlon. http://www.nature.com/nphys/journal/vaop/nprelaunch/full/nphys128.html
"The inner radiation belt extends from a height of 600 km to 6,000 km above the Earth's equator. It is filled mainly with the byproducts of collisions as cosmic rays hit the Earth's atmosphere. The outer radiation belt is much larger and more variable in extent, but on average lies between 12,000 and 60,000 km in altitude"
Propulsion Technologies and orbital change
The following section looks at
methods of raising the altitude of a satellite.
My thanks to Achim DH2VA and Flario who suggested the Hohmann transfer method
and to Bob N4HY for the reference to the NASA site.
Additionally, Achim has produced an Excel spreadsheet which calculates delta V
and orbital change from fuel mass ratio etc.
To change from a lower orbit (A) to a higher orbit (C), an engine is first fired in the opposite direction from the direction the vehicle is traveling. This will add velocity to the vehicle causing its trajectory to become an elliptic orbit (B). This elliptic orbit is carefully designed to reach the desired final altitude of the higher orbit (C). In this way the elliptic orbit or transfer orbit is tangent to both the original orbit (A) and the final orbit (C). This is why a Hohmann transfer is fuel efficient. When the target altitude is reached the engine is fired in the same manner as before but this time the added velocity is planned such that the elliptic transfer orbit is circularized at the new altitude of orbit (C).
Following the 'publication' of the MEOSAT idea, Achim DH2VA / HB9DUN
responded with an excellent Excel spreadsheet programme which calculates the
required delta V to raise the height of an orbit.
Here is Achims AMSATBB posting with the URL and notes.
http://gulp.physik.unizh.ch/meosat_propulsion.xls
download the excel sheet. I started from the following boundary
conditions:
starting orbit circular (no GTO.. this complicates things).. something a
Russian launcher can deliver (if the staging works well :( )
Target orbit circular.. transfer orbit Hohmann type, so you have two
burns: one to change the low orbit into an elliptical with apogee at the
height of the high orbit and a second burn to circularize. This is the
theoretically most efficient transfer.
The input fields are in blue (start height, target height, Isp, dry
mass, fueled mass) and the results (required delta_V, achievable
delta_v) are in red. I filled in some values from David's website
including the Isp for the bipropellant engine and a 50/50 partitioning
of fuel and dry mass for a 20 kg launch mass. It is just possible..
Please play around and report errors.. I do not take responsibility for
any wrong mission planning :)
73s Achim, DH2VA/HB9DUN
Propulsion for Dummies
I'll be the first to admit
that I don't know a whole lot about propulsion, but having conceived the MEOSAT
idea, I thought I should investigate the practicalities of getting a nanosat /
microsat from LEO to a more interesting higher altitude orbit.
We all know a little about propulsion systems that have been used on some
satellites. AO10, AO13 and AO40 all used a 400 Newton engine fueled by
Hydrazine and Nitrogen Tetroxide. AO40 of course weighed 500kg and so
needed more fuel than its smaller predecessors. The fuels used on the phase 3
satellites were very hazadous and the use of such dangerous materials also means
the control systems to operate a bipropellant engine are complex. Even the
experts can get this wrong. So what other methods are there for generating
thrust for a small satellite? The answer appears to include:
1) Cold gas thrusters.  As used on AO40 (Ammonia) or SSTL's SNAP1
mission (32.6 grams of butane)
2) Hot gas thrusters  e.g. The ARC Jet using Ammonia and an electrical ignition
circuit. As fitted, but not used on AO40. Or the electrolysis of water into
Hydrogen and Oxygen which can then be burned in a combustion chamber.
3) Chemical single burn engines.  e.g The solid rocket boosters used on the
shuttle / Ariane 5. Or a much smaller version.
4) Electric or Ion propulsion  As used on the moon orbiter SMART1
Any engine will have a bewildering list of characteristics. A good site
linking these is. http://www.grc.nasa.gov/WWW/K12/airplane/specimp.html
This presents the various formulae and shows
how they are derived.. The math's looks a little intimidating at the top,
but gets easier as you progress. If you don't fancy a weekend with a thumping
headache lets see if I can distill just enough of the essentials, with a couple
of extreme examples, to do some basic orbital transfer calculations using
Achim's spreadsheet. To calculate the potential increase in altitude from a
particular propulsion system we will need to know the mass of the satellite with
fuel. The mass without fuel, hence the mass of the fuel and finally we need to
know the efficiency of the propulsion system which is known as the Specific
Impulse ( Isp )
Thrust: Any engine will be designed to produce a certain amount of thrust. This is a measure of the force exerted in Newtons when the engine is operating. The thrust can vary from very small. 0.050 Newtons for SNAP1 to 19600 Newtons for the Russian Fregat engine. The engine will also be designed to provide this amount of thrust for a particular period of time. 297 seconds for SNAP1. 877 seconds for Fregat.
Total Impulse: This is the product of thrust and time. e.g. 0.050 x 297 = 14.85 Newton Seconds for SNAP1. Over 17 million NS for Fregat
Fuel mass: This is simply the mass of fuel. 32.6 grams or 0.0326kg for SNAP1. 5350kg for the Fregat
Mass
flow rate:
This is the rate that the fuel flows through the engine. 32.6g / 297
seconds = 0.0001097643 grams per second for SNAP1.
For the ariane, the fuel consumption is rather more at 6.1kg per
second
Specific Impulse: (Isp) This is a very important parameter and it is a measure of the efficiency of a propulsion system..
Specific Impulse = Thrust / mass flow rate / gravitational
acceleration
Where Thrust is in Newtons. Mass flow rate in
kg/second Gravitational acceleration is a constant at 9.81
metres per second.
Example 1. For
SNAP1 Isp = 0.05 / ( 0.0326 / 297 ) /
9.81 = 46
s  The unit
of Isp is the second  Low efficiency (cold gas)
Example 2 For Russian Fregat . Isp = 19600 N / 6.1kg/s / 9.81m/s = 327 s  This is typical for a bipropellant rocket.  high efficiency
The following comparison of propellant technologies has been copied from the European Space Agency site and gives comparative figures for chemical and electric propulsion systems. Note that the Ion engine can only produce very small level of thrust but the Isp is very high indicating that it is much more efficient than a bipropellant engine. Overall the Ion drive uses fuel more efficiently and would appear to be a good choice for deep space flight, but only if flight time wasn't an issue.
Comparison of propulsion technologies 

Chemical 
Electric 

Small monopropellant thruster  Fregat Main Engine 
SMART1 Hall Effect Thruster 

Propellant  Hydrazine  Nitrogen tetroxide / Unsymmetrical dimethyl hydrazine  Xenon 
Specific 
200  320  1640 
Thrust (N)  1  1.96 x 10^{4}  6.80 x 10^{2} 
Thrust 
1.66 x 10^{5}  877  1.80 x 10^{7} 
Thrust 
46  0.24  5000 
Propellant 
52  5350  80 
Total 
1.1 x 10^{5}  1.72 x 10^{7}  1.2 x 10^{6} 
Fregat produces ~ 14 times the total impulse of SMART1's engine, but uses nearly 70 times more propellant mass to do so. The hydrazine thruster produces less than a tenth as much total impulse while using 65% of the propellant mass. 
So, after that diversion into propellants lets
get back to MEOSAT and raising the altitude.
Let's start with cold gas propulsion. The original specification for the
SSTL butane system was for an ISP of 60. If we use that figure we can look at
the potential for cold gas.
Example. Cold Gas Propulsion
Satellite mass empty 1.8kg
Fuel mass 1kg
Isp 60
Run Achim's spreadsheet and enter a starting height of 800km and a destination
altitude of 900km.
Enter 60 in the column for ISP and the start mass of 2.8kg and the final mass of
1.8kg.
Now look at the results. You can see that to raise the
orbit by 100km the required total change in velocity or delta V is 51 metres per
second.
Can our cold gas system produce this velocity change? Yes
!
Actually, the programme shows that our 1kg of propellant can achieve a total
delta V of 260 m/s.
So, how high can we go? Just increase the desired height while
looking at the required delta V. When the required delta V equals 260m/s,
that's how high we can go.
How high is that? Well for the above example the maximum achievable
altitude is 1328km. That is a total increase of 528km for 1kg of propellant.
Not Medium Earth Orbit, but a much more interesting LEO.
Example. Solid fuel propulsion
For the second example let's try using a solid fuel.. To get the data on a
simple solid rocket motor I used the performance details presented for model
rocket engines. These are intended to lift model rockets a few hundred feet into
the air but the manufacturers present all of the data we need. Thrust, Total
impulse. Thrust duration and mass of propellant. The result is an Isp or
Specific thrust of 240.
(ESTES E98. Total Impulse 30NS duration 2.8 seconds mass 35.8g )
I believe it would be possible to increase the
efficiency so lets use 280.
Of course, the solid fuel would only fire once, and we need two burns for a
Hohmann transfer but lets calculate it anyway....
Enter the information as before. 1kg of fuel, 1.8kg final mass Isp 280. Start at 800km destination altitude 1300km
Notice that now we have a more efficient propulsion system (higher Isp) the maximum achievable delta V has increased over the cold gas system. It was 260m/s, but with the solid fuel it's now 1214m/s.
How high will this take us? Can we get to the safe zone above 7000km? Can we use more fuel / better fuel or technology?
There are many variations to try.......Just
enter the details into the spreadsheet and have some fun
If you
find a winning combination let us all know.......Over to you................