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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 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.
https://en.wikipedia.org/wiki/Passive_radar#Typical_illuminators

 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
  

Tuesday, October 4, 2016

A smaller, faster version of the SpaceX Interplanetary Transport System to Mars, UPDATED, 10/15/2016.

Copyright 2016 Robert Clark


 In the blog post "A SpaceX Heavy Lift Methane Rocket, Page 2", I proposed some architectures for a Mars transport rocket. This was based on a quite large 1,800,000 lb. vacuum thrust of the methane-powered Raptor.

 However, recently Elon Musk discussed the current version of the SpaceX Mars transport model called the Interplanetary Transport System (ITS). Here they are going back to a smaller version of the Raptor, at ca. 660,000 lb. vacuum thrust. In this version however, their booster will be quite large at ca. 7,000 metric tons (mT) gross weight.

 Because of the reduced size of the Raptor this will require 42 engines on the booster. However, interestingly the size of the upper stage will be similar to the size of the booster I discussed in "A SpaceX Heavy Lift Methane Rocket, Page 2". So you could get a Mars launch booster by using this upper stage instead as a booster.



But because of the smaller engines in the SpaceX formulation they will use nine of the Raptors on the upper stage. I stated in the earlier blog post I wanted to use at most 5 of the larger Raptor engines to emulate the safety record of the 5 large engines on the Saturn V booster. However, SpaceX seems to have gotten the 9 engines on the Falcon 9 to work, and in any case you could just use the booster to send the cargo and habitats to space and use high safety rockets to launch the crew to meet up with the transport craft in space.

 The objection could be made however, that this is supposed to be just an upper stage, not a booster stage. However, at his IAC presentation of this Mars transport system he stated that the upper stage in both the spaceship and tanker form could be SSTO. Furthermore, the tanker he said could be used a fast intercontinental transport craft. This means necessarily they would have to be able to launch from the ground. So it is not too much of stretch to assume they could be used as a first stage.



An advantage of making this smaller upper stage the actual booster is that Elon has said they will have a development craft within 4 years. So if we make now a smaller upper stage to go with it, we could have a valid Mars transport craft at that early date. If we made this new upper stage correspondingly 1/3rd size, then we would be able to get 1/3rd the crew size to Mars, so a crew of ca. 35 to Mars.

 However, interestingly we might be able to use already existing upper stages on existing rockets for the upper stage, for instance possibly the famous Centaur upper stage used on the Atlas V or the Ariane 5 core itself used here as an upper stage.

 We can estimate how much we could get to LEO using the ITS tanker as the booster and the Ariane 5 core as the upper stage. The required delta-v to LEO is 30,000 ft/sec about 9,100 m/s:

Modern Engineering for Design of Liquid-Propellant Rocket Engines, p.12

 Since we can get a high 465 s vacuum Isp for a hydrolox engine just by using a nozzle extension we'll assume that value for the Ariane 5 engine. Also we'll assume we can get a vacuum 382 s Isp for the ITS tanker by using altitude compensation.

 Then for the propellant and dry mass values for the ITS tanker and Ariane 5 core we could get 225 metric tons to LEO:

382*9.81ln(1 + 2500/(90 + 170 + 225)) + 465*9.81ln(1 + 158/(12 + 225)) = 9,140 m/s.

 This makes clear another key advantage of this architecture: whereas the original SpaceX ITS would require five flights of the ITS to refuel the upper stage spaceship, with this smaller version a single flight would be able to carry the spaceship to orbit as well as its fuel for its flight towards Mars.

 But what would be the crew size for this smaller upper stage? We can estimate it by making a comparison to the delta-v possible in accordance with the stats of the ITS spaceship:

382*9.81ln(1 + 1950/(150 + 450)) = 5,400 m/s.

 So we want the Ariane 5 case to be able to reach a delta-v of 5,400 m/s when fully refueled and firing in space headed towards Mars:

465*9.81ln(1 +158/(12 +55)) = 5,500 m/s.

 So this is a payload of about 55 metric tons. This is about 1/8th the mass for the ITS case, so we can estimate the crew size to be 1/8th also, so to a crew of 12.

  Elon in his IAC presentation says the ITS carrying its 100 member crew might be able to reach Mars in 80 days at a particular close Mars opposition. This is dependent on the departure delta-v however. In the blog post "Propellant depots for interplanetary flight". I noted that at a higher departure delta-v possible by using a smaller 6 metric ton habitat for only a crew of 3, the Ariane 5 used as the in space propulsion stage might be able to make it to Mars in only 35 days, when leaving at such a particularly close Mars opposition.


  Bob Clark

UPDATE, 10/15/2016:

 Dr. John Schilling's launch performance calculator is back up. This allows us to produce a more accurate payload estimate. The vacuum thrust for the 382 Isp Raptor is 3.5 meganewtons, 3,500 kN. So 9 would be 31,500 kN. We'll also increase the vacuum thrust of the Vulcain engine on the Ariane 5 to 1,450 kN in accordance with the increased vacuum Isp of 465 s. Inputting these and the other 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:  177767 kg
95% Confidence Interval:  150063 - 210818 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 should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.



 So the calculator estimates 178 metric tons. This is less than the 225 metric ton estimate using just the rocket equation, but it still means a single flight could carry enough payload to fully refuel an Ariane 5 core upper stage for a flight to Mars.

  Bob Clark

Monday, June 27, 2016

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

Copyright 2016 Robert Clark
(patents pending)


 Some calculations show a surprising increase in the amount of payload that can be carried by a single-stage-to-orbit rocket (SSTO) by using altitude compensation [1], such as the aerospike, even multiple times more than possible without it. Indeed, the calculations revealed that for an already high propellant fraction stage such as the Falcon 9, alt. comp. gives the SSTO a better cost per kilo ratio than the two stage rocket (!) 

 This was a surprising result since during much of the era of orbital rockets it was received wisdom that SSTO's were not technically feasible. Then, it gradually became accepted it could be done, but it was then felt it would not be worthwhile because of the small payload. Therefore it is quite remarkable that the exact opposite of this is true, the SSTO is more cost effective than the TSTO (two-stage-to-orbit) when using altitude compensation [1]. 

 But the usefulness of altitude compensation is not just for SSTO's. The payload for a two-stage to orbit launcher can be increased 25% by using it [2]. And triple-cored rockets such as the Delta IV Heavy, and Falcon Heavy can have their payload doubled when using altitude compensation in concert with cross-feed fueling [2]. Moreover, by using alt. comp., simple pressure-fed stages that are within the technical means of most university engineering departments can be made to make suborbital [3] and orbital launchers [4].

 However, an argument has been made that transforming already existing engines to altitude compensation such as the aerospike would be expensive since it would require changing the combustion chamber to a toroidal shape. Then I investigated other means of achieving altitude compensation other than the aerospike [5].

 One of these methods was to use high temperature carbon nanotube "rubber" [6] as a nozzle extension. This could be attached to the nozzle of already existing engine nozzles and be variably extended as the rocket gained altitude.

 But could we use metals for this purpose? The metal would have to be stretchable as is rubber to become twice as long or more as the nozzle is extended. Normally though metal can only be stretched by a fraction of its original length before fracturing and even then it takes quite a large amount of force to do the stretching.

 There is a scenario though where metals can be stretched for a longer length and at a small amount of required force, that is at elevated temperatures. This is through forging. This takes place while the metal is still solid. The forging temperature [7] is where the metal is more malleable but below the melting temperature. It is commonly in the range of 60% of the melting temperature. Then the idea would be as the nozzle becomes heated as the engine is firing it would become more and more easily extended further out. 

 For how to extend, that is stretch, the nozzle, one possibility would be to use high pressure inert gas such as helium injected within the hollow walls of the nozzle to stretch it you as would for blowing up a a hollow balloon. Another would be actuators attached to the end to stretch it out.

 For either method you would want the nozzle to maintain the usual bell nozzle shape. You could have the wall thickness vary along the nozzle's length so that as it is stretched out the required shape is maintained. You might also have ribs along the vertical length of the nozzle to help encourage the stretching to proceed in the desired direction.

 Another consideration is that you don't want the nozzle to reach a degree of heating so that it reaches the melting point. An interesting fact about rocket nozzles and combustion chambers is that they actually operate at temperatures above the melting point of the metal composing them. The reason why they don't melt is that for a material to undergo the phase change from solid to liquid, not only does the temperature have to be at the melting point, but a sufficient quantity of heat dependent on the material has to be supplied to the material, the enthalpy of fusion [8].

 Then rocket engines have cooling mechanisms applied to the chamber and nozzle walls to draw away the heat supplied by the combustion products so that this amount of heat is never applied to chamber and nozzle. One key method that is used for high performance engines is regenerative cooling. This is where the fuel is circulated through channels in the walls of the engine to draw away the heat.

 Another factor to limit the temperature and heat applied to the nozzle is that this is envisioned as an attachment to a usual, static nozzle. However, as the engine exhaust is expanded out by a bell nozzle the temperature drops. So for the attachment at the bottom of the usual nozzle, the temperatures it would have to withstand would be reduced.

 A diagram showing the stress-strain curve at elevated temperatures for titanium alloys is here [9]:


  The strain at room temperature is commonly only a fraction of a percent, ca. 0.2%, or 0.002. But here at elevated temperatures in the range of 800C to 1,050C, we see the strain can reach .7, and likely above with continued pressure applied.



REFERENCES.

1.)Thursday, November 7, 2013
The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2.

2.)Monday, January 11, 2016
Altitude Compensation Improves Payload for All Launchers.

3.)Thursday, January 15, 2015
NASA Technology Transfer for suborbital and air-launched orbital launchers.

4.)Thursday, August 13, 2015
Orbital rockets are now easy.
http://exoscientist.blogspot.com/2015/08/orbital-rockets-are-now-easy.html

5.)Saturday, October 25, 2014
Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

6.)Carbon Nanotube Rubber Stays Rubbery in Extreme Temperatures.
Liming Dai
Angew. Chem. Int. Ed. 2011, 50, 4744 – 4746
http://case.edu/cse/eche/daigroup/Journal%20Articles/2011/Dai-2011-Carbon%20Nanotube%20Rubb.pdf

7.)Forging temperature.
https://en.wikipedia.org/wiki/Forging_temperature

8.)Enthalpy of Fusion.
https://en.wikipedia.org/wiki/Enthalpy_of_fusion

9.)MODELLING HIGH TEMPERATURE FLOW STRESS CURVES OF
TITANIUM ALLOYS
Z. Guo, N. Saunders, J.P. Schillé, A.P. Miodownik
Sente Software Ltd, Surrey Technology Centre, Guildford, GU2 7YG, U.K
http://www.sentesoftware.co.uk/media/2524/flow_stress_curve.pdf



Tuesday, April 12, 2016

Combined amateur telescopes for asteroid detection.

Copyright 2016 Robert Clark

 NASA is conducting an interesting program to get the public involved in the upcoming ORISIS-REx mission to retrieve a sample from an asteroid. It is asking amateur astronomers to make observations of known asteroids using their telescopes:

Target Asteroids!

 However, a slight modification of this program should allow it to also to discover unknown asteroids. This article discusses that even an 8-inch scope equipped with a CCD camera can discover new asteroids:

Hunting Asteroids From Your Backyard
By: Dennis Di Cicco | July 28, 2006
There are no hard and fast rules regarding the telescope or CCD camera needed for asteroid work. To be effective, the system should record stars as faint as 18th magnitude with a single, 4-minute exposure. Almost any CCD camera on an 8-inch telescope can do this under a clear, dark sky.
http://www.skyandtelescope.com/observing/celestial-objects-to-watch/hunting-asteroids-from-your-backyard/

 The article discusses down to magnitude 18. But combining the observations of many of these scopes acting in concert should allow the discovery of asteroids of weaker magnitude and therefore smaller size.

 As discussed in the article, CCD's can have imaging artifacts where a pixel will show as lit but it's not really corresponding to a light photon hitting the device. Moreover, the weaker the imaging source, the more difficult it is distinguish these imaging artifacts from a real light source.

 However, since these imaging artifacts are occurring at random, the idea would be to have several of the amateur scopes from different parts of the world focused on the same spot in the sky. Then several of the scopes' CCD's registering a hit on a pixel corresponding to the same point in the sky at the same time would be taken as indicating a real light source.

 The scopes would have to have a high degree of sky location specificity and timing synchronization for this to work.
 
 Another aspect of the imaging artifacts of the CCD's is that at low imaging illumination the CCD might correctly register a lit pixel but at a later time not register it. For individual scopes used to detect asteroids, it's done by noticing the light source moving between exposures. But if the imaging light source is too weak the CCD for the scope might not register the light source the second time to detect the motion. Then in this proposal of using multiple scopes, you also need to be able to correlate a second detection by another scope as indicating the light source moved, even if this scope did not detect the light source the first time. All the information would need to be correlated at a central site for this to work.

 Then after sufficient numbers of scopes give a high level of confidence the asteroid is indeed there, larger professional telescopes could be used to confirm the detection.

 This would have importance also for planetary protection purposes since it would allow the detection of smaller asteroids.

Credit and Financial Rewards for the Discovery?
 Certainly the amateur astronomers whose scopes detected the asteroid should get credit for the discovery. But an intriguing question of financial rewards arises because of the companies such as Planetary Resources, Inc. and Deep Space Industries that are working towards returning valuable minerals from asteroids. According to this article an asteroid potentially worth $5 trillion in platinum passed nearby to Earth last year:

‘Platinum’ asteroid potentially worth $5.4 trillion to pass Earth on Sunday.
Published time: 18 Jul, 2015 11:21
https://www.rt.com/news/310170-platinum-asteroid-2011-uw-158/

 There are very many near Earth asteroids still to be discovered. Then one can imagine these coordinated amateur scopes detecting one of these highly valuable asteroids. If one of them is eventually used to recover valuable minerals should the amateur astronomers who discovered it take part in the financial rewards?

 Not an easy question but it is notable that it would increase the interest and participation of amateur astronomers in the program. In view of its potential importance for planetary defense purposes this participation should be encouraged.

   Bob Clark