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SpaceX BFR tanker as an SSTO.



 
 
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  #1  
Old October 28th 17, 05:40 AM posted to sci.space.policy
William Mook[_2_]
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Posts: 3,840
Default SpaceX BFR tanker as an SSTO.

On Monday, October 23, 2017 at 2:30:22 PM UTC+13, Sylvia Else wrote:
On 23/10/2017 1:37 AM, Robert Clark wrote:
The SpaceX BFR tanker can serve as a reusable SSTO by switching to a
winged, horizontal landing mode:


On a quick scan through, it looks to me as if he's ignored the mass of
the landing gear, which in a typical aircraft is 10% of the empty weight.

A non-air breathing SSTO ends up with a horribly small payload fraction,
which means in turn that only a modest increase in vehicle mass can wipe
out the payload completely.

Sylvia.


With VTOL using rockets, and with an advanced 'active' landing platform, no landing gear is needed. Merely hold down clamps.

Catching Balls with Robots

https://www.youtube.com/watch?v=R6pPwP3s7s4

Which I discussed here;

Active Landing Pads - catch rockets

https://vimeo.com/37102557

BFR Mass Budget:

BFR system consists of 85 t second stage with 1,250 t propellant capacity - delivering 150 t propellant for transfer to a BFR spaceship - after burning 1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR spaceship.

85 t / 1250 t = 6.8% structure fraction.

Way more than needed.

We have 9.75 t of propellant to bring the 85 t structure to a soft landing from 0.34 km/sec terminal velocity.

10% of the 95 t structure + propellant is 9.5 t - set that aside for landing gear.

Tank fraction is 3.5% of the original capacity - 43.75 t.

7 Raptor engines at 1 t each is 7 t. That's 50.75 t sub-total. 9.5 t landing gear 60.25 t sub-total.

Add 9.75 t propellant - and that's 70.00 t.

Add 15 t thermal protection and other hardware - and consumables - and you have your 85.00 t budget.


* * *


https://www.youtube.com/watch?v=tdUX3ypDVwI

Specifications

Complete................ BFR booster........... BFR spaceship
LEO Payload........... reusable: 150 t (330,000 lb)
................................ expendable: 250 t (550,000 lb)
Return Payload....... 50 t (110,000 lb)
Diameter................. 9 m (30 ft)
Length................... 106 m (348 ft)....... 58 m (190 ft) 48 m (157 ft)
Maximum mass...... 4,400 t (9,700,000 lb) 1,335 t (2,943,000 lb)
Engines.................. 31× SL Raptors 3× SL + 4× vacuum Raptors[37]
Thrust.................... 52.7 MN (11,800,000 lbf) 12.7 MN (1,300 tf; 2,900,000 lbf) total
Propellant Capacity 1,100 t (2,400,000 lb): 240 t CH4 + 860 t O2
Empty weight......... 85 t (187,000 lb)
Ads
  #2  
Old October 28th 17, 03:44 PM posted to sci.space.policy
Robert Clark[_5_]
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Posts: 245
Default SpaceX BFR tanker as an SSTO.



With VTOL using rockets, and with an advanced 'active' landing platform, no
landing gear is needed. Merely hold down clamps.

Catching Balls with Robots

https://www.youtube.com/watch?v=R6pPwP3s7s4

Which I discussed here;

Active Landing Pads - catch rockets

https://vimeo.com/37102557

BFR Mass Budget:

BFR system consists of 85 t second stage with 1,250 t propellant capacity -
delivering 150 t propellant for transfer to a BFR spaceship - after burning
1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR
spaceship.

85 t / 1250 t = 6.8% structure fraction.

Way more than needed.

We have 9.75 t of propellant to bring the 85 t structure to a soft landing
from 0.34 km/sec terminal velocity.

10% of the 95 t structure + propellant is 9.5 t - set that aside for
landing gear.

Tank fraction is 3.5% of the original capacity - 43.75 t.

7 Raptor engines at 1 t each is 7 t. That's 50.75 t sub-total. 9.5 t
landing gear 60.25 t sub-total.

Add 9.75 t propellant - and that's 70.00 t.

Add 15 t thermal protection and other hardware - and consumables - and you
have your 85.00 t budget.



Keep in mind there is a difference between the tanker and the spaceship
versions of the upper stage. The spaceship version has higher dry mass
because it has the passenger quarters and supplies for a 6 month trip to
Mars carrying 100 colonists. Since the tanker version doesn't carry this
passenger compartment it has much lighter dry mass. This is illustrated by
the ITS upper stage introduced last year:

https://i.imgur.com/GsyREf7.png

The ITS spaceship had a dry mass of 150 tons, and the ITS tanker a dry mass
of 90 tons.
The half-size BFR spaceship has a dry mass of 85 tons. Then applying similar
scaling to the BFR tanker we can estimate its dry mass as 50 tons.

It is the BFR tanker that would be used to deliver payload to LEO.

Bob Clark

  #3  
Old October 28th 17, 11:36 PM posted to sci.space.policy
William Mook[_2_]
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Posts: 3,840
Default SpaceX BFR tanker as an SSTO.

You can buy them and do anything you can get a license to do. So, have at it!

With a 3.2 km/sec exhaust speed and the requirement to boost through a 9.2 km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec lost due to air drag and gravity losses) we have a propellant fraction of;

u = 1 - 1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36%

So, with a 3.5% structure fraction you could place 2.14% into LEO.

With a TSTO-RLV and the same structure fraction, dividing the delta vee by 2 to make each stage 4.6 km/sec we have;

u = 1 - 1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25%

So with a 3.5% structure fraction you could push 20.25% of each stage through 4.6 km/sec.

So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~ 4.1%

Nearly double the take off weight into orbit.

Let's do three stages! 9.2 km/sec / 3 = 3.067 so,

u = 1 - 1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65%

So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85% payload on each stage.

0.348531572876311 cubed is 4.23%

A slight, but measureable improvement.

So, you can see there is a significant benefit to two stages over one.

Now, if your structure fraction is higher, the benefit of staging is higher. If we go from 3.5% to 5.0% structure fraction, we have

1 stage to orbit -0.64% - 155.85 t per t on orbit
2 stage to orbit - 3.51% - 28.44 t per t on orbit
3 stage to orbit - 3.71% - 26.96 t per t on orbit


On Sunday, October 29, 2017 at 3:44:43 AM UTC+13, Robert Clark wrote:

With VTOL using rockets, and with an advanced 'active' landing platform, no
landing gear is needed. Merely hold down clamps.

Catching Balls with Robots

https://www.youtube.com/watch?v=R6pPwP3s7s4

Which I discussed here;

Active Landing Pads - catch rockets

https://vimeo.com/37102557

BFR Mass Budget:

BFR system consists of 85 t second stage with 1,250 t propellant capacity -
delivering 150 t propellant for transfer to a BFR spaceship - after burning
1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR
spaceship.

85 t / 1250 t = 6.8% structure fraction.

Way more than needed.

We have 9.75 t of propellant to bring the 85 t structure to a soft landing
from 0.34 km/sec terminal velocity.

10% of the 95 t structure + propellant is 9.5 t - set that aside for
landing gear.

Tank fraction is 3.5% of the original capacity - 43.75 t.

7 Raptor engines at 1 t each is 7 t. That's 50.75 t sub-total. 9.5 t
landing gear 60.25 t sub-total.

Add 9.75 t propellant - and that's 70.00 t.

Add 15 t thermal protection and other hardware - and consumables - and you
have your 85.00 t budget.



Keep in mind there is a difference between the tanker and the spaceship
versions of the upper stage. The spaceship version has higher dry mass
because it has the passenger quarters and supplies for a 6 month trip to
Mars carrying 100 colonists. Since the tanker version doesn't carry this
passenger compartment it has much lighter dry mass. This is illustrated by
the ITS upper stage introduced last year:

https://i.imgur.com/GsyREf7.png

The ITS spaceship had a dry mass of 150 tons, and the ITS tanker a dry mass
of 90 tons.
The half-size BFR spaceship has a dry mass of 85 tons. Then applying similar
scaling to the BFR tanker we can estimate its dry mass as 50 tons.

It is the BFR tanker that would be used to deliver payload to LEO.

Bob Clark

  #4  
Old October 31st 17, 06:55 AM posted to sci.space.policy,sci.physics,rec.arts.sf.science,sci.astro
Robert Clark[_5_]
external usenet poster
 
Posts: 245
Default SpaceX BFR tanker as an SSTO.

"William Mook" wrote in message
...
You can buy them and do anything you can get a license to do. So, have at
it!
With a 3.2 km/sec exhaust speed and the requirement to boost through a 9.2
km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec
lost due to air drag and gravity losses) we have a propellant fraction of;
u = 1 - 1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36%
So, with a 3.5% structure fraction you could place 2.14% into LEO.
With a TSTO-RLV and the same structure fraction, dividing the delta vee by
2 to make each stage 4.6 km/sec we have;
u = 1 - 1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25%
So with a 3.5% structure fraction you could push 20.25% of each stage
through 4.6 km/sec.
So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~
4.1%
Nearly double the take off weight into orbit.
Let's do three stages! 9.2 km/sec / 3 = 3.067 so,
u = 1 - 1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65%
So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85%
payload on each stage.
0.348531572876311 cubed is 4.23%
A slight, but measureable improvement.
So, you can see there is a significant benefit to two stages over one.
Now, if your structure fraction is higher, the benefit of staging is
higher. If we go from 3.5% to 5.0% structure fraction, we have

1 stage to orbit -0.64% - 155.85 t per t on orbit
2 stage to orbit - 3.51% - 28.44 t per t on orbit
3 stage to orbit - 3.71% - 26.96 t per t on orbit


Thanks for that calculation. It is correct, for the most part. There is the
fact that structure fraction only refers to the rocket stage itself without
payload. But when you calculate the fraction that reaches orbit, that's of
the entire mass including payload. So it's not quite accurate to subtract
off the structure fraction from this. But since the payload is a small
percentage of the entire stage mass it's a small discrepancy.

However, a key consideration that should be taken into account is that to
optimize the payload for a SSTO you really should use altitude compensation.
Your structure fraction for the stage of 3.5% is very good, but the exhaust
speed of 3.2 km/sec isn't very good for a vehicle you want to be SSTO.
Altitude compensation allows you to maximize your vacuum Isp while
optimizing your sea level thrust at launch as well.

Since we're discussing in the context of the Raptor engine I'll use
estimates for methane engines. I'll use a rocket engine analysis program to
estimate the possible vacuum Isp with methane fuel:

http://www.propulsion-analysis.com/index.htm

The specs of the Raptor in its latest incarnation are given he

https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg

If you use the cited combustion chamber pressure of 250 bar of the Raptor,
but give it an expansion area ratio of 300, possible with altitude
compensation, then the vacuum Isp can be in the range of 390 s according to
the rocket engine analysis program.

In contrast, the sea level Raptors to be used on the BFR will have only a
vacuum Isp of 356 s, and even the vacuum optimized Raptors to be used only
have a vacuum Isp of 375 s. By using altitude compensation you don't have to
make these trades of how many sea level vs. vacuum engines to use. All the
engines will have the optimal performance both at sea level and at vacuum
with altitude compensation.

So let's redo your calculation assuming the good structure fraction of 3.5%
but using altitude compensation to improve the Isp to 390 s, exhaust speed
of 3,824 m/s.

Then the proportion of the rocket, dry mass plus payload, that reaches orbit
is: 1/exp(9200/3,824) = .09019, then subtracting off the .035 for structure
fraction, the payload fraction it would be .09019 - .035 = .05519, about
5.5%, significantly better than the 2.14% you get with only a 3.2 km/s
exhaust speed. Even more notable is that the payload fraction of the SSTO
with altitude compensation is even better than that of the TSTO without it.

But to compare apples to apples, let's calculate the payload fraction
assuming both stages of the TSTO get the 390 s Isp. In that case,
1/exp(4600/3,824) = .300313 is the final stage mass at burn out, dry mass
plus payload. So subtracting off the structure fraction, the payload
fraction for each stage would be .300313-.035 = .265313. Then squaring this
because it is a two-stage, the payload fraction would be .07039, about 7%,
well above the payload fraction of the TSTO without altitude compensation.

This shows altitude compensation can make a significant improvement even for
multistage vehicles. The payload fraction is still better than the SSTO
case, but it is not as radically better as in the case without altitude
compensation. When you take into account you don't have the extra expense of
the second stage and of integrating the two stages together, the SSTO can be
useful for launching smaller payloads.


Bob Clark
--
----------------------------------------------------------------------------------------------------------------------------------
Carbon nanotubes can revolutionize 21st-century technology IF they can be
made arbitrarily long while maintaining their strength.
Some proposals to accomplish that he
From Nanoscale to Macroscale: Applications of Nanotechnology to Production
of Bulk Ultra-Strong Materials.
American Journal of Nanomaterials.
Vol. 4, No. 2, 2016, pp 39-43. doi: 10.12691/ajn-4-2-2 | Research Article.
http://pubs.sciepub.com/ajn/4/2/2/
----------------------------------------------------------------------------------------------------------------------------------

  #5  
Old November 2nd 17, 07:28 AM posted to sci.space.policy
William Mook[_2_]
external usenet poster
 
Posts: 3,840
Default SpaceX BFR tanker as an SSTO.

SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO

The S-II stage of the Saturn V.

5 J2 engines 1,000,000 lbs thrust
Take off Weight 1,060,000 lbs weight
Structure 80,560 lbs
Ve=9,235 mph.

This can carry 34,100 lbs into LEO. About the size of a Gemini Capsule!

Today, we could do a Dragon Capsule!

Total take off weight with payload: 1,094,100 lbs.

Now, an aerospike engine, has a higher thrust to weight than a bell nozzle. It also has altitude compensating features which you've described below.

https://www.youtube.com/watch?v=-0Y0FS8Z1Qk

So, if you take 7 J2 pumpsets, and outfit a toroidal aerospike engine around it at the base of the S-11, and equip the zero height aerospike engines with a heat shield, you should using modern materials, attain the same structural fraction as was achieved in the 1960s with the S-II whilst increasing thrust to 1,400,000 lbs.

With a 9,235 mph exhaust speed a thrust of 1,400,000 lbf requires 3,323.8 lbs per second of propellant flow.

Acceleration is 42.2 ft/sec/sec minus 32.2 ft/sec/sec is 10.0 ft/sec/sec.

If we accelerate straight up for 100 seconds a total of 332,380 lbs of propellant have been burned. At that point the vehicle masses 727,620 lbs. So, acceleration rises at constant thrust to 1,400,000 lbf / 727,620 lbs = 62.2 ft/sec/sec minust 30 ft/sec/sec net that's

a = 42.2*exp(0.00388 * t)
V = integral(42.2*exp(0.00388*t) -32.2 ) dt, t=0 to 100 = 1,935.68 ft/sec = 1,319.8 mph.
D = integral(10,876.3 * exp(0.00388*t) - 32.2*t - 10876.3) dt, t=0 to 100 = 80,156.1 ft = 15.18 miles

Of course, you could start tilting the rocket, and calculate the cosine of the angle as vertical, and the sine of the angle as horizontal add them together to get total velocity, take the ratio to get the tangent and so forth..

av=cos(theta*t)*(42.2*exp(0.00388*t)-32.2)
ah=sin(theta*t)*(42.2*exp(0.00388*t))

at=sqrt(av^2+ah^2)
V=integral(sqrt(av^2+ah^2))
D=integral(integral(sqrt(av^2+ah^2)))

At $1,000 per pound, construction cost, the entire system would cost $85 million for the booster $35 million for the payload - $120 million total.


Reused 1200 times - the cost is $100,000 per flight. Propellant is $500,000 in quantity and maintenace $82,000 - between flights. That's $20 per pound.

http://www.spacefuture.com/archive/h..._systems.shtml


* * * *

On Tuesday, October 31, 2017 at 7:55:12 PM UTC+13, Robert Clark wrote:
"William Mook" wrote in message
...
You can buy them and do anything you can get a license to do. So, have at
it!
With a 3.2 km/sec exhaust speed and the requirement to boost through a 9..2
km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec
lost due to air drag and gravity losses) we have a propellant fraction of;
u = 1 - 1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36%
So, with a 3.5% structure fraction you could place 2.14% into LEO.
With a TSTO-RLV and the same structure fraction, dividing the delta vee by
2 to make each stage 4.6 km/sec we have;
u = 1 - 1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25%
So with a 3.5% structure fraction you could push 20.25% of each stage
through 4.6 km/sec.
So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~
4.1%
Nearly double the take off weight into orbit.
Let's do three stages! 9.2 km/sec / 3 = 3.067 so,
u = 1 - 1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65%
So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85%
payload on each stage.
0.348531572876311 cubed is 4.23%
A slight, but measureable improvement.
So, you can see there is a significant benefit to two stages over one.
Now, if your structure fraction is higher, the benefit of staging is
higher. If we go from 3.5% to 5.0% structure fraction, we have

1 stage to orbit -0.64% - 155.85 t per t on orbit
2 stage to orbit - 3.51% - 28.44 t per t on orbit
3 stage to orbit - 3.71% - 26.96 t per t on orbit


Thanks for that calculation. It is correct, for the most part. There is the
fact that structure fraction only refers to the rocket stage itself without
payload. But when you calculate the fraction that reaches orbit, that's of
the entire mass including payload. So it's not quite accurate to subtract
off the structure fraction from this. But since the payload is a small
percentage of the entire stage mass it's a small discrepancy.

However, a key consideration that should be taken into account is that to
optimize the payload for a SSTO you really should use altitude compensation.
Your structure fraction for the stage of 3.5% is very good, but the exhaust
speed of 3.2 km/sec isn't very good for a vehicle you want to be SSTO.
Altitude compensation allows you to maximize your vacuum Isp while
optimizing your sea level thrust at launch as well.

Since we're discussing in the context of the Raptor engine I'll use
estimates for methane engines. I'll use a rocket engine analysis program to
estimate the possible vacuum Isp with methane fuel:

http://www.propulsion-analysis.com/index.htm

The specs of the Raptor in its latest incarnation are given he

https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg

If you use the cited combustion chamber pressure of 250 bar of the Raptor,
but give it an expansion area ratio of 300, possible with altitude
compensation, then the vacuum Isp can be in the range of 390 s according to
the rocket engine analysis program.

In contrast, the sea level Raptors to be used on the BFR will have only a
vacuum Isp of 356 s, and even the vacuum optimized Raptors to be used only
have a vacuum Isp of 375 s. By using altitude compensation you don't have to
make these trades of how many sea level vs. vacuum engines to use. All the
engines will have the optimal performance both at sea level and at vacuum
with altitude compensation.

So let's redo your calculation assuming the good structure fraction of 3.5%
but using altitude compensation to improve the Isp to 390 s, exhaust speed
of 3,824 m/s.

Then the proportion of the rocket, dry mass plus payload, that reaches orbit
is: 1/exp(9200/3,824) = .09019, then subtracting off the .035 for structure
fraction, the payload fraction it would be .09019 - .035 = .05519, about
5.5%, significantly better than the 2.14% you get with only a 3.2 km/s
exhaust speed. Even more notable is that the payload fraction of the SSTO
with altitude compensation is even better than that of the TSTO without it.

  #6  
Old November 2nd 17, 03:56 PM posted to sci.space.policy
Greg \(Strider\) Moore
external usenet poster
 
Posts: 624
Default SpaceX BFR tanker as an SSTO.

"William Mook" wrote in message
...

SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO

The S-II stage of the Saturn V.


Stop, please just stop.
If you really want to post this stuff, start your own thread. Please.
Stop hijacking other threads.

--
Greg D. Moore http://greenmountainsoftware.wordpress.com/
CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net
IT Disaster Response -
https://www.amazon.com/Disaster-Resp...dp/1484221834/

  #7  
Old November 2nd 17, 08:43 PM posted to sci.space.policy
Fred J. McCall[_3_]
external usenet poster
 
Posts: 9,764
Default SpaceX BFR tanker as an SSTO.

"Greg \(Strider\) Moore" wrote:

"William Mook" wrote in message
...

SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO

The S-II stage of the Saturn V.


Stop, please just stop.
If you really want to post this stuff, start your own thread. Please.
Stop hijacking other threads.


He can't stop Mookjacking threads. I don't think he knows how to
start his own.


--
You are
What you do
When it counts.
  #8  
Old November 5th 17, 04:48 PM posted to sci.space.policy,sci.physics,rec.arts.sf.science,sci.astro
Robert Clark[_5_]
external usenet poster
 
Posts: 245
Default SpaceX BFR tanker as an SSTO.


Since we're discussing in the context of the Raptor engine I'll use
estimates for methane engines. I'll use a rocket engine analysis program to
estimate the possible vacuum Isp with methane fuel:

http://www.propulsion-analysis.com/index.htm

The specs of the Raptor in its latest incarnation are given he

https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg

If you use the cited combustion chamber pressure of 250 bar of the Raptor,
but give it an expansion area ratio of 300, possible with altitude
compensation, then the vacuum Isp can be in the range of 390 s according to
the rocket engine analysis program.


That image of the Raptor on freelists.org site is not available. It's
available he

http://spaceflight101.com/spx/wp-con...17-Musk-19.jpg

Bob Clark

--
----------------------------------------------------------------------------------------------------------------------------------
Carbon nanotubes can revolutionize 21st-century technology IF they can be
made arbitrarily long while maintaining their strength.
Some proposals to accomplish that he
From Nanoscale to Macroscale: Applications of Nanotechnology to Production
of Bulk Ultra-Strong Materials.
American Journal of Nanomaterials.
Vol. 4, No. 2, 2016, pp 39-43. doi: 10.12691/ajn-4-2-2 | Research Article.
http://pubs.sciepub.com/ajn/4/2/2/
----------------------------------------------------------------------------------------------------------------------------------

  #9  
Old November 6th 17, 11:23 AM posted to sci.space.policy,sci.physics,rec.arts.sf.science,sci.astro
Jeff Findley[_6_]
external usenet poster
 
Posts: 1,706
Default SpaceX BFR tanker as an SSTO.

In article ,
says...


Since we're discussing in the context of the Raptor engine I'll use
estimates for methane engines. I'll use a rocket engine analysis program to
estimate the possible vacuum Isp with methane fuel:

http://www.propulsion-analysis.com/index.htm

The specs of the Raptor in its latest incarnation are given he

https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg

If you use the cited combustion chamber pressure of 250 bar of the Raptor,
but give it an expansion area ratio of 300, possible with altitude
compensation, then the vacuum Isp can be in the range of 390 s according to
the rocket engine analysis program.


That image of the Raptor on freelists.org site is not available. It's
available he

http://spaceflight101.com/spx/wp-con...17-Musk-19.jpg


Sure, it's theoretically possible to do such things. But, SpaceX has
actually been somewhat conservative on how they've approached engine
design. Altitude compensation would add a new facet of "bleeding edge
technology" which would introduce more risk into their program
(especially schedule risk since this is a relatively unknown area beyond
past ground testing).

That and they've long ago decided that a fully reusable TSTO is the way
to go. If they're successful at landing the BFR first stage on the
launch platform, that would eliminate *a lot* of work to get it ready to
fly again. They should be able to inspect it, stack the BFR upper stage
on top, refuel it, and fly it again.

SpaceX hasn't been successful due to using bleeding edge technology.
They've been successful by avoiding bleeding edge technology.

The one thing they did embrace, because they had to, was supersonic
retro-propulsion with liquid fueled rocket engines. They only did that
because their initial "simpler" plan, which was parachute recovery of
first stages in the ocean, quite simply didn't pan out.

Jeff
--
All opinions posted by me on Usenet News are mine, and mine alone.
These posts do not reflect the opinions of my family, friends,
employer, or any organization that I am a member of.
 




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