Tuesday, August 22, 2017

Orbital rockets are now easy, page 2: solid-rockets for cube-sats.

Copyright 2017 Robert Clark

  In the blog post "Orbital rockets are now easy", I argued that with altitude compensation, liquid-fueled orbital rockets become within the capabilities of most university undergraduate labs.

 Here I'll argue solid-fuel rockets can also be built by amateurs to reach orbital space.

 I was impressed by this university teams launch to 144,000 feet of a suborbital solid-fuel rocket:

USC Rocket Propulsion Laboratory Breaks Record.
Amy Blumenthal | March 16, 2017
Student-run RPL launches rocket of own design to 144,000 feet.

  However, using essentially "off-the-shelf" components, you can get a 3-stage solid rocket of comparable size to actually reach orbit. Moreover, I was surprised to see after running a launch simulation program that you don't even need altitude compensation to accomplish it.

 By essentially off-the shelf, you could use solid-rocket motors sold to amateurs but with an addition of a lightweight carbon-composite casing that amateurs such as the USC team already have been making themselves for their own rockets.

 You would have Cesaroni Technology, or other solid motor manufacturer, make their motors with only a thin aluminum case, not meant to hold the full combustion chamber pressures. The amateurs would then make the carbon composite case rated for the full combustion chamber pressure.

 Cesaroni also makes commercial sounding rockets with carbon composite casings but this is likely to be more expensive than if the amateurs make it themselves.

 This motor by Cesaroni Technology, or similar motors by other manufacturers, could form the basis of the orbital rocket:

  This motor though only has a mass ratio of about 2.5 to 1, adequate for amateurs doing high power rocketry test flights, but not for an orbital rocket. However, you can save about half of the weight of the aluminum casing used by replacing it by carbon composite.

 Cesaroni was able to do this for the suborbital rocket SpaceLoft they constructed for UP Aerospace, Inc.:

CTI rocket motor successfully powers the launch carrying the ashes of astronaut and James Doohan - April 30, 2007.
On April 28th, a Spaceloft™XL rocket successfully completed a round-trip space flight launched from Spaceport America. This rocket was developed by UP Aerospace Inc. of Hartford, Conn. The rocket carried a wide variety of experiments and payloads, which included the cremated remains of Star Trek's "Scotty", James Doohan and NASA astronaut and pioneer Gordon Cooper. In addition, the cremated remains of more than 200 people from all walks of life were onboard. Also flown into space on the SL-2 Mission were dozens of student experiments from elementary schools to high schools to universities - from across America and worldwide - as well as innovative commercial payloads.
The flight was a successful demonstration of the rocket motor developed and built by Cesaroni Technology Incoporated (CTI). CTI started the design process in September of 2005. CTI specializes in low cost propulsion systems for military and space applications and used its experience to develop an affordable, reliable propulsion system for the rocket. The motor has a carbon fiber composite case and a monolithic solid propellant grain that is bonded to the casing.
Watch the post-launch coverage as carried by local television station KRQE here (9 Mb)
Watch the launch as carried by the BBC here (1 Mb)
Watch pre-launch coverage as carried by CTV Toronto here (9 Mb)
Watch pre-launch coverage as carried by CBC Toronto here (9 Mb)
Technical data for the UPA-264-C rocket motor

 So we'll assume the Cesaroni Pro150 can get a 0.8 propellant fraction by using carbon composite casing. We'll round off the propellant mass to 20 kg, and take the dry mass as 5 kg. We'll take this as the third stage of our rocket.

 For the lower stages, the size of stages commonly are in the range of 3 to 5 times larger than the succeeding stage. We'll take the second stage as 4 times larger than the third stage at a 80 kg propellant weight and 20 kg dry weight. For the first we'll take it as larger by an additional factor of 4 to a 320 kg propellant weight and 80 kg dry weight.

 Now for the specific impulses for the stages. This commercial solid motor has a similar sea level Isp as the Cesaroni solid rocket motor Pro150:

Star 37.
Thiokol solid rocket engine. Total impulse 161,512 kgf-sec. Motor propellant mass fraction 0.899. First flight 1963. Solid propellant rocket stage. Burner II was a launch vehicle upper stage developed by Boeing for the Air Force Space Systems Division. It was the first solid-fuel upper stage with full control and guidance capability developed for general space applications.
AKA: Burner 2;TE-M-364-1. Status: First flight 1963. Number: 180 . Thrust: 43.50 kN (9,779 lbf). Gross mass: 621 kg (1,369 lb). Unfuelled mass: 63 kg (138 lb). Specific impulse: 260 s. Specific impulse sea level: 220 s. Burn time: 42 s. Height: 0.84 m (2.75 ft). Diameter: 0.66 m (2.16 ft).
Thrust (sl): 33.600 kN (7,554 lbf). Thrust (sl): 3,428 kgf.

 So we'll estimate the vacuum Isp of the Cesaroni Pro150 to be in the Star 37's range of 260 s. However, rocket stages can get higher vacuum Isp's by using longer nozzles. A 285 s Isp is not uncommon for solid rockets motors with vacuum optimized nozzles, such as the Star 48.

 So we'll take the Isp for the second and third stage as 285 s.

 For the thrust, we'll take the thrust of the third stage as the same as the Pro150 of 8 N, and assume the thrust for the second and first stage scale according to their size so to 32 N and 128 N, respectively.

 Now use Dr. John Schilling's launch performance calculator to estimate the payload.

 The input page looks like this:

 Note there some quirks of this program you need to be aware of if you use it. First, always use the vacuum values for the Isp's and thrust numbers, since the program already takes into account the diminution at sea level. Second, always set the "Restartable Upper Stage" option to "No", rather than the default "Yes", otherwise the payload will be reduced. Third, always set the launch inclination to match the launch site latitude, otherwise the payload will be reduced. This is related to the fact that changing the orbital plane involves a delta-v cost. So for the Cape Canaveral launch site, the launch inclination should be set to 28.5 degrees.

Now, here's the result:

 So a payload to LEO of 7 kg. And this with standard nozzles, no altitude compensation required.

  To save costs, I'm envisioning making the components as much "off-the-shelf" as possible. But among its standard products Cesaroni offers the Pro150 as the largest motor. So to get the larger second and first stages, we would have to combine multiple copies of this motor.

  I could cluster them in parallel, but for the first stage that would be 16 of them, and you would have the problem of simultaneous ignition with that many motors.

  So what I'm envisioning is take 4 copies of the Pro150 stacked vertically one on top of the other for the second stage, then cluster 4 of these second stage motors in parallel for the first stage.

 The question is though about the vertical stacking is how much the thrust scales in this case. If for the solid motors the propellant burned from the bottom upwards, then the thrust would be the same as for a single motor, you would just get 4 times longer burn time.

  But that's not how large solid motors work. Actually, they have a hollow region in the center so the propellant burns from the inside surface, proceeding outwards. In this case, you have a greater amount of propellant being burned per second because of the larger vertical surface area with the stacked segments. In fact, the thrust scales linearly with the number of segments.

 By the way, the reason why I don't just also stack the first stage vertically, is because of the thinness of the rocket that would result. The Pro150 is about 3 feet long and 1/2 foot wide. If you stacked vertically 16 for the first stage, 4 for the second, and 1 for the third, that would be a rocket 63 feet high but only 1/2 foot wide, for a ratio of length to width of over 120 to 1.

 This ratio of length to width is called the "fineness ratio". Rocket engineers don't like for it to be higher than about 20 to 1 because of the severe bending loads that would result. The upgraded version of the Falcon 9 has been noted for its long, skinny profile, and has a fineness ratio of about 20 to 1. The Scout solid rocket had a fineness ratio of about 24 to 1. Solid rockets can support a higher fineness ratio because their thicker walls can withstand higher loads. Still, 120 to 1 would very likely be too high.

  So to avoid this I decided to form the first stage by clustering in parallel four copies of the second stage. Note here these four clustered motors arranged around the second stage will provide the full thrust for the first stage while the central second stage motor will not fire until the four clustered motors are jettisoned.

 It would be possible though to get a single vertically stacked motor using multiple segments if the segments were shorter, resulting in a smaller rocket. For instance, there is a market for cubesats at only 1 kg mass to orbit. If you made the solid motor segments only about 1/2 foot long by cutting the Pro150 into 6 segments, you could take one of these smaller segments as the third stage, 4 segments for the second stage, and 16 segments for the third stage.

 The Cesaroni Pro150 retails for about $3,000 and in general the Cesaroni solids cost in the range of $100 per kg of the motor mass:

Cesaroni O8000 White Thunder Rocket Motor.   
Product Information
Brandname:  Pro150 40960O8000-P                  Manufacturer:  Cesaroni Technology
Man. Designation:  40960O8000-P                    CAR Designation:  40960 O8000-P
Test Date:  4/10/2008                                   
Single-Use/Reload/Hybrid:  Reloadable             Motor Dimensions mm:  161.00 x 957.00 mm (6.34 x 37.68 in)
Loaded Weight:  32672.00 g (1143.52 oz)         Total Impulse:  40960.00 Ns (9216.00 lb/s)
Propellant Weight:  18610.00 g (651.35 oz)       Maximum Thrust:  8605.10 N (1936.15 lb)
Burnout Weight:  13478.00 g (471.73 oz)          Avg Thrust:  8034.50 N (1807.76 lb)
Delays Tested:  Plugged                                      ISP:  224.40 s
Samples per second:  1000                                  Burntime:  5.12 s

 So take the cost of the third stage, derived from the Cesaroni Pro1050, as $3,000, and the second stage 4 times larger as $12,000, and the first stage larger by an additional factor of 4 as $48,000. So $63,000 for a smallsat launcher with a 7 kg payload to orbit.

 Several universities have created their cubesats and smallsats to be launched piggyback on large rockets such as the Falcon 9. However, the solid rocket launcher formed from essentially off the shelf components could be built by any interested university itself thus creating their own launcher and satellite.

 Despite the small size of such satellites, and their low construction cost, the launch cost is still not cheap when sent piggyback. SpaceX for their latest incarnation of the Falcon 9 is charging about $60 million for a 20,000 kg payload to LEO, about $3,000 per kilo. But the price is much higher than that for small payloads that have to ride piggyback on launchers. For instance Spaceflight Industries charges about $100,000 per kilo to book such flights. But a 1 kg cubesat launch would only cost in range of $9,000 for one of these dedicated solid-rocket launchers.

 The remaining entrants to the Google Lunar X-Prize will have to pay expensive launch costs for their spacecraft to the Moon. But with the university teams using their own solid rocket launchers, the launch costs would be so cheap the teams could afford to make many attempts to win the lucrative $30 million prize to soft-land on the Moon, and many more teams could have remained in the race.

 Also, both DARPA and the Army funded programs to develop such small dedicated launchers (liquid fueled), with their ALASA and SWORDS programs. But both their programs failed. They wanted to get about 25kg to 50kg to orbit for a launch cost of $1,000,000, about $20,000 to $40,000 per kilo. But the small size solid rockets will be able to beat this price, moreover will be more operationally responsive by using solid rockets. In a follow up post I'll discuss the payload can be more than doubled by using altitude compensation reducing the per kilo cost even further.

National security implications.
 Recently, there has been some discussion on creating Ultra Low Cost Access to Space (ULCATS). See for instance this study:


 Most of the discussion has been about how this would improve U.S. capabilities. But surprisingly little has been about what are the national security implications of any university world-wide and many knowledgeable amateur groups world-wide launching their own rockets to orbit.

 To prepare for this, which will be here like tomorrow, this discussion must begin now.

    Bob Clark

Note: thanks to former aerospace engineer and math professor GW Johnson for helpful discussion on this topic on the forum:

Amateur solid-fueled rockets to *orbital* space?

Tuesday, August 1, 2017

Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.

Copyright 2017 Robert Clark

 In prior posts I gave some possibilities for achieving altitude compensation, [1],[2], [3], [4].The importance of this is they increase the payload both for single stage and multistage rockets.

 Another possibility is illuminated by this:

 Only it would use pressurize gas rather than popcorn to expand out the nozzle.

   Bob Clark


1.)Altitude compensation attachments for standard rocket engines, and applications.

2.)Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.

3.)Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.

4.)Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.

Sunday, May 7, 2017

Test flights of the Falcon Heavy for missions to the moons of Earth and Mars, Page 1.

Copyright 2017 Robert Clark

  The SpaceX Red Dragon lander mission to Mars on the Falcon Heavy has been pushed back to 2020, perhaps to return a Mars surface sample. SpaceX though plans for two Falcon Heavy test flights for the latter part of this year, 2017.

  Elon has discussed testing recovery of the stages on these first test flights which will reduce payload. He has also discussed putting a "fun" payload on one of them, like his cheese wheel on the first Dragon test flight.

  I suggest instead missions be undertaken of great scientific and practical importance, missions to the moons of Earth and Mars. 

Flight to a Permanently Shadowed Crater on the Moon.
 Abundant evidence suggest ice water deposits in the permanently shadowed craters on the Moon. This has been proposed to be used to produce orbital propellant depots. This would radically reduce the mass that would need to be launched to orbit for a Mars mission since most of this mass is just propellant. 

 There have also been some tantalizing indications from the LCROSS mission of valuable metals in the shadowed craters. Then the first space mining missions may be to the Moon rather than the asteroids.
 I'll estimate the delta-v to land on the Moon using this diagram:

 The delta-v to GTO (geosynchronous transfer orbit) is 2.5 km/sec. Then after that according to the delta-v diagram we need an additional 3.2 km/sec to land on the Moon. We wish to use the cargo version of the Dragon to land on the Moon. This weighs about 5.5 metric tons (mT) fueled with its own propellant, while the FH can get 26.7 mT to GTO. 

 So the idea would be to get extra delta-v by using the smaller mass of the Dragon capsule. To estimate this we'll need the specs for the upper stage of the Falcon Heavy, same as for the Falcon 9's upper stage, 348 s Isp for the Merlin 1D FT, approx. 107.5 mT propellant load, and approx. 4 mT dry mass. Then the delta-V this upper stage achieves with the 26.7 mT payload is 348*9.81*Ln(1 + 107.5/(4 + 26.7)) = 5,136 m/s.

 So by reducing the payload mass from 26.7 mT to 5.5 mT we want this upper stage to achieve a delta-v of:

5.1(to reach GTO) + 3.2(to land on the Moon) = 8.3 km/sec .

 And calculating the delta-v of the stage with the reduced payload we get:

348*9.81*Ln(1 + 107.5/(4 + 5.5)) = 8,572 m/s.

 This is above the needed 8.3 km/s, though close. Actually we'll get somewhat better than this because the lower stages having to loft a lighter payload will be able to provide more delta-v than before.

 Also, we actually will use the Draco thrusters on the Dragon to do the actual landing since the FH upper stage would put the capsule to high up if it were to land vertically, and it's thrust is so high achieving the stable landing is made difficult.

 That raises another difficulty because of the low thrust of the Draco thrusters. There are 18 Dracos of the cargo Dragon each of thrust level 400 N, for a total of 7,200 N. This can lift 7,200/9.81 =  734 kg in Earth's gravity. In the lunar gravity at 1/6th g, the Dracos could lift, 4,404 kg. But the fueled mass is 5,500 kg.

 There are a couple of things we can do to lighten the Dragon. We could remove both the parachute and thermal protection systems since the capsule won't be returning to Earth in this mission.  These weigh about 5% each of the landed mass, so about 10% all together. So this shaves 420 kg off the landed mass.

 Another possibility would be to replace the Dracos with Superdracos, which have many times greater thrust. But I'm not sure how well these would fit in the same housing for the Dracos.

 We could also remove most of the pressure vessel for landing on the airless Moon. From images of the Dragon's pressure vessel, this could be a significant mass:

 For the rover, we might use a copy of the Mars Pathfinder mission. NASA often makes two or more copies of its spacecraft for testing purposes. Then we could use one of these copies. This weighed only 264 kg for the lander plus 10.5 kg for the Sojourner rover.

 Other possibilities for a lightweight rover might be those being developed independently by entrants to the Google Lunar X-prize:

Flight to Phobos, the mysterious moon of Mars.
 A great scientific mystery also is the make-up of the Mars moon Phobos. Flyby missions showed it to have surprisingly low density. Serious scientific speculation included that it may actually be hollow. Current theories are though that it may be analogous to a "rubble-pile" type asteroid. This is not known for sure however. A lander mission may help to resolve the issue.

 Note also that key to Elon's plan for manned flights to Mars is getting the fuel for the return trip from Mars. Taking the fuel from the Martian moons instead would have advantages such as low gravity for getting the fuel to an orbiting propellant depot. Then these first flights to the Martian moons could serve as scout missions for water ice deposits.

  The Falcon Heavy test flights this year will be outside the optimal launch window in 2018. This means they will require higher delta-v to reach Mars, and higher delta-v to slow down on reaching the destination. This limits the mass that can be transported to and landed on Mars, in addition to the expense of the extra in-space stages required.

 Then I will suggest here a method that has long been proposed for arriving at Mars but never attempted, aerocapture. This slows down a craft arriving at Mars by plunging deep within the atmosphere so that minimal propellant burn is required. Note, that if these tests missions using aerocapture succeed then this will suggest it will work to solve the problem of landing large masses on Mars such as a crew habitat, a key enabling technology for manned flights to Mars.

 For the delta-v required to depart from Earth I'll use the orbital calculation program:

Trajectory Planner.

 This provides the delta-v's required for the Hohmann tranfer orbits between the various planets. The program provides pork-chop plots that allow you to estimate departure and arrival delta-v's dependent on departure time.

 The program though uses Modified Julian Date format, which can be converted to standard date format here:

 For a Dec. 23, 2017 departure, which is given in Modified Julian Date format of 58110 in the "Trajectory Planner", the delta-v Hohmann transfer delta-v is 6.155 km/s. We then need to calculate the delta-v needed on leaving Earth orbit. On the Orbiter-Forum discussion forum for the Orbiter space simulation program this formula was provided by member Dgatsoulis:

 \Delta V = \sqrt{V_{\infty}^2 + V_{esc}^2} - V_{orb}

where V_{\infty} is the hyperbolic excess velocity (departure deltaV from trajectory planner).

V_{esc} is the local escape velocity, aka the escape velocity for the parking orbit altitude.

V_{esc} = \sqrt{\frac{2GM_{planet}}{R_{planet}+alt}}

where G is the gravitational constant, M_{planet} is the planet's mass, R_{planet} is the planet's radius and alt is the altitude of the parking orbit.

V_{orb} is the parking orbit velocity.

V_{orb} = \frac{V_{esc}}{\sqrt{2}} 

 Same applies for arrival. If you want to simply calculate the periapsis velocity and not the orbit insertion/injection dV, then don't use the V_{orb} term.


 So the  delta-v on leaving Earth orbit is:
 This 1.11 km/sec more than the usual delta-v to make a Trans Mars Injection during the optimal departure windows of 3.8 km/s.

 We need to calculate how much mass the FH upper stage could get to this higher delta-v of 4.89 km/s. By the FH specs it can get 16.8 mT to Trans Mars Injection.This FH upper stage with the 16.8 mT payload mass can do 348*9.81Ln(1 + 107.5/(4 + 16.8)) =  6,211 m/s delta-v. So with a smaller mass we want to achieve 6,211 + 1,110 = 7,321 m/s delta-v. This can be done with a 10 mT payload:

338*9.81Ln(1 + 107.5/(4 + 10)) = 7,377 m/s.

 Now we have to calculate how much is the speed on arrival at Mars.  The Trajectory Planner gives the "arrival" speed as 4.275 km/s. However, again this is not the speed the spacecraft would have on entering Mars's atmosphere. This is instead the speed at which it arrives at Mars's position in its orbit around the Sun, i.e., the Hohmann orbit delta-v needed to be supplied to match Mars' solar orbital speed.

 To get the entry speed into Mars' atmosphere, use the Dgatsoulis formula above without the Vorb term. Using 5.0 km/s as the escape velocity for Mars we get:

 If all we wanted was to slow down to enter Mars orbit then we would subtract off from this by aerocapture to bring the speed down to Mars' orbital velocity of 3.56 km/s. However we also want to be put it on a trans Phobos insertion from Mars. By the delta-v chart above we need an additional .9 km/s, so to 4.46 km/s. So by the aerocapture we only need to slow it down by about 2.11 km/s.

 This should be well within the capabilities of aerocapture. However, the payload mass may be as high as 10 mT. The question is could the dragon's approx. 10 square meter base provide sufficient air drag to slow down that high mass, and would its heat shield be thick enough?

 In follow up posts I'll present some preliminary calculations that suggest that plunging deep into Mars atmosphere, skimming the tree-tops so to speak, should allow large masses such as this to be slowed at such high entry speeds.

 With the payload of the FH as high a 10 mT, the rovers and equipment that could be transported could be 4.5 mT above the 5.5 mT fueled weight of the cargo Dragon. But according to the delta-v chart we still need 0.5 km/s delta-v to land on Phobos. The cargo Dragon has a delta-v capability of about 600 m/s with its Draco thrusters for the Dragon capsule alone. So this should be sufficient, but it would not be if the extra cargo was several metric tons. So we could keep the cargo low as for a Mars Pathfinder sized rover or we could add additional propellant tanks to increase the landing capability.

 The possible cargo carried by the Dragon being as high as 4.5 mT suggests though we should try to make use of that cargo space. One possibility would be the processing equipment to produce ISRU (in situ resource utilization) propellant. Perhaps a rocket to do a sample return. Possibly orbiting imaging spacecraft for Phobos or Mars. Others?

    Bob Clark

Note: thanks to members of Keithth G and DGatsoulis for helpful discussions on this topic and member Piper, for writing the Trajectory Planner program.

Tuesday, April 25, 2017

About the launch abort system for the New Shepard capsule.

Copyright 2017 Robert Clark

 Blue Origin has revealed the format of its suborbital tourism capsule for the New Shepard suborbital launcher:

Take a Peek Inside Blue Origin’s New Shepard Crew Capsule.
Published: 29 Mar , 2017
by Nancy Atkinson

   The cylinder in the middle is the launch abort motor. It is only supposed to fire in case of an emergency to pull the capsule away from the rocket launcher.

 Normally, it would not even fire. Still its presence inside the passenger cabin is rather disconcerting. Moreover, it is a solid rocket motor. For solid motors, the combustion chamber is the entire rocket, so if a failure, i.e., a breech does occur it can happen anywhere along the motors length.

 A Blue Origins video animation from 2015 shows the solid rocket escape motor with handholds at about the 2:25 point:

 Be careful to mind your head while floating though!

   The reason Blue Origin decided to put the abort motor inside the cabin likely was for reasons of positioning of the center of gravity(CG) with respect to the center of pressure(CP). A well known rocket stability rule of thumb is the center of pressure should be below the center of gravity

The trunk and fins helped that for the SpaceX launch abort test by bringing the CP rearward:

 But compare this to the Blue Origin abort test:

  Notice that the capsule is gyrating while the rocket motor is firing. This would be very unpleasant for the passengers since they would be subjected to high g's while being thrown right and left, albeit while strapped in.

 Then for these reasons I suggest giving the New Shepard a trunk with fins as has the SpaceX Dragon capsule.

 This could be done by instead of having the ring structure at the top of the New Shepard stay attached to the New Shepard, let it act as the trunk for the capsule:

 Then you would move the solid rocket abort motor down into this structure, so it is no longer inside the passenger compartment.

 However, this ring structure does have a function as far as the landing of the New Shepard rocket; it holds the fins and the speed brakes used during the landing:

 So how could we maintain those functions if that ring structure is instead attached to the capsule? Two possible approaches you could duplicate it so the New Shepard has its own as does the capsule. 

 Or another possibility would be to have the ring structure only detach along with the capsule only during an abort scenario. For the normal launch, with no abort, the ring structure would stay attached to the New Shepard rocket, carrying also inside the abort motor, while the capsule detaches for the normal flight to suborbital space.

 But if there is a need for an abort, the solid rocket abort motor would fire carrying the ring structure and the capsule away from the New Shepard. In this scenario where there would need to be an abort presumably there would be a failure of the New Shepard anyway and you would not expect to recover it.

  Bob Clark

Wednesday, March 15, 2017

Satellite dishes and satellite phones for radio astronomy and passive radar detection.

Copyright 2017 Robert Clark

Asteroid Detection.
 In the blog post "Combined amateur telescopes for asteroid detection", I suggested using multiple small amateur telescopes in concert to act as a giant astronomical instrument to make dim observations in the optical range. Could we do the same with multiple satellite dishes or satellite phones to make dim radio observations? 

 There is a technique called "passive radar" that uses reflected radio waves from aircraft that originate from surrounding radio transmissions such as from television and radio stations:

Passive Radar.
3. Typical illuminators
Passive radar systems have been developed that exploit the following sources of illumination:
Analog television signals
FM radio signals
Cellular phone base stations
Digital audio broadcasting
Digital video broadcasting
Terrestrial High-definition television transmitters in North America
GPS satellites (GPS reflectometry).
Satellite signals have generally been found to be inadequate for passive radar use: either because the powers are too low, or because the orbits of the satellites are such that illumination is too infrequent. The possible exception to this is the exploitation of satellite-based radar and satellite radio systems. In 2011, researchers Barott and Butka from Embry-Riddle Aeronautical University announced results claiming success using XM Radio to detect aircraft with a low-cost ground station.

 The difficulty in using satellite transmissions for the detections previously is that they just use a single ground station for the reception of the reflected signals. Instead of this, suppose we used millions of satellite dishes or radios or satellite phones to make the detections?

 As with the case of multiple amateur telescopes, you couldn't form a coherent signal from this method. But like in the optical case you could make correlations from which you could make a probabilistic estimate of the likelihood of an actual detection.

 There is an additional difficulty however. We are envisioning using satellites at geosynchronous orbit, about 35,000 kilometers out in space. We would detect asteroids closer than this distance by their blocking the satellite signals from being detected by satellite dishes or phones.

 However, the asteroids would tend to direct the reflected signals back out to space rather than towards the Earth, except for the case where the asteroid is along a line from the satellite towards the limb of the Earth, and with the dishes/phones along the limb. But this would be relatively few asteroids and dishes/phones so precisely placed in the right position.

 So in actuality for this method to work we would be looking for holes, deletions, in the signal. Such deletions in the satellite signal would be small for each dish or phone. But by correlating the signals of millions of them we can determine statistically that it represents a real detection.

 This would only be for detecting asteroids rather close in, since they would be inside the distance of geosynchronous orbit. This would still be useful since from multiple observations we could determine their orbits. And such asteroids that came so close in would have a higher probability of presenting an impact hazed on a future orbital pass.

 But could we also detect asteroids further out? Some proportion of the signal from the GEO satellites likely escapes past the sides of the Earth to proceed to the other side. And this proportion of the signal likely is increased by the signals bouncing off the ionosphere. Then these signals could proceed further outwards to be reflected back to Earth by more distance asteroids.

 The strength of the signal leaking past Earth would be reduced so the reflected signals would also be reduced. But in this case you are making actual positive detections rather than looking for holes in the signal so all in all the results could be just as effective as in the close in asteroid case.

Aircraft detection.
 A problem with detecting aircraft on intercontinental flights is that when they fly over the oceans they fly too far from the radar stations on land to be detected. Then perhaps the method of satellite signal detections by multiple dishes/phones can be used to track such aircraft as well. This may give a us a method to finally locate the missing airliner Malaysian Airlines Flight 370. The flight was lost three years ago but there may have been some satellite TV, radio or phone customers who saved programs or phone conversations at that time for which the recorded digital data can be reviewed to reconstruct a detection of the aircraft.

  Bob Clark

Tuesday, March 14, 2017

A smaller, faster version of the SpaceX Interplanetary Transport System to Mars, Page 2: triple cores for larger payloads.

 Copyright 2017 Robert Clark

 In the blog post "A smaller, faster version of the SpaceX Interplanetary Transport System to Mars", I suggested using just the upper stage of the ITS to get a booster for a Mars rocket, using an existing Ariane 5 core as an upper stage. This would be much cheaper and faster than the 7,000 metric ton, 42 engine booster that SpaceX was planning.

 Elon Musk says SpaceX plans to have the smaller upper stage built by 2020. So we could possibly have a Mars transport system by then since the Ariane 5 as an upper stage already exists. However, by using triple cores of the ITS upper stage we could also get a system of the larger size SpaceX is proposing.

 We'll input the data into Dr. John Schillings payload estimation program. In the calculator, select "No" for the "Restartable Upper Stage" option, rather than the default "Yes", otherwise the payload will be reduced. Select Cape Canaveral as the launch site, and input 28.5 degrees for the launch inclination to match the latitude of Cape Canaveral, otherwise the payload will be reduced.

 We'll also use the 382 s Isp of the vacuum version of the Raptor. Altitude compensation allows even engines used on first stage boosters to have the same vacuum Isp as upper stages engines.

 We'll use also crossfeed fueling. As I have argued before this is a well-known technique having been used for decades on jet airliners. To emulate crossfeed fueling with the Schilling calculator, enter in 2/3rds the actual propellant load in the field for the sideboosters, and enter in (1 + 2/3) times the actual propellant load in for the first stage propellant load.

 So in the side boosters propellant field enter in (2/3) * 2,500,000 kg = 1,667,000 kg. And in the first stage propellant field enter in (1 + 2/3) * 2,500,000 = 4,167,000 kg.

 For the thrust fields, enter in the vacuum thrust for 9 vacuum Raptors, since the calculator always takes as input the vacuum values, even for first stages and side boosters. The vacuum thrust for the 382 Isp vacuum Raptor is 3.5 meganewtons, 3,500 kN. So 9 would be 31,500 kN. Enter in also the vacuum Isp 382 s.

 For the second stage, we'll increase the vacuum thrust of the Vulcain engine on the Ariane 5 to 1,450 kN in accordance with an increased vacuum Isp of 465 s, since we can get this higher vacuum Isp by just using a nozzle extension. For the dry mass input 12,000 kg and propellant 158,000 kg. Inputting these specs in the calculator results in:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  504575 kg
95% Confidence Interval:  426107 - 597674 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and shou
ld not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.

 This is comparable to the payload mass of the expendable version of SpaceX's ITS. This would save greatly on development costs when not having to develop the larger booster. The launch cost would also be greatly reduced since judging by the Falcon Heavy, using triple cores only increased the price 50% over that of the single core rocket.

    Bob Clark

Sunday, October 30, 2016

Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.

Copyright 2016 Robert Clark
(patent pending)

 On the blog page "Altitude compensation attachments for standard rocket engines, and applications", I suggested various methods to accomplish altitude compensation with already existing engines. One method was a sort-of "inverted aerospike". It consisted of a movable spike pointed inward, rather than pointing outward as with the standard aerospike:

 There are two disadvantages to this method. First the spike has to be movable so that adds mechanical complexity. Secondly, the size of the outer, fixed nozzle in order to achieve high Isp at high altitude has to be large. But this nozzle will be used all the way from the ground, so this will induce high drag at low altitude.

 The reason why this nozzle has to be large is because you are not really using the altitude compensating capacity of a shaped spike on exit from the nozzle. The only purpose of the movable spike is to vary the size of the exit plane of the nozzle, to provide a variable area ratio.

 But could we use a fixed nozzle and the usual outward-pointing aerospike? This would have the advantages that we could use the altitude compensating capacity of the usual aerospike, so we could use a shorter nozzle, and also have a fixed spike, reducing mechanical complexity. 

 The problem with this with a usual engine is you would need to change to a toroidal combustion chamber, an expensive change to an engine. So instead of this, we will also use an inward pointing spike so that the exit of the nozzle has a toroidal shape:

  This now has two advantages. We will be using this as an attachment to a usual ground-firing engine and nozzle. Since these already expand the exit gases to a certain extent, you would need a much shorter, slimmer and lighter outward-pointing spike to accomplish the rest of the expansion at high altitudes. The usual aerospike has to accomplish the full expansion from ca. 100 bar combustion chamber pressures to near vacuum pressures at high altitude, requiring a large and heavy spike.

  Another advantage is that nozzles for sea-level-firing engines actually overexpand the exit gases at sea level. This is because you want a longer nozzle to achieve at least moderate performance also at high altitude. But now, with the addition of the inward-pointing spike you can reduce the pressure at exit of the nozzle to that of sea level by reduction of the exit plane area. This will also improve the performance at sea level.

   Bob Clark