William Mook
2009-12-19 21:50:19 UTC
Andy wrote;
> > Three steerable re-usable SRBs producing 12.5 million newtons, EACH,
> > of thrust would lift the shuttle to orbit without its 26.5 tonne tank,
William wrote;
> No it wouldn't.
William notes: (my explanation of the rocket equation is deleted)
Andy replies:
> ======================================
>
> "Each engine can generate almost 1.8 meganewtons (MN) or 400,000 lbf of
> thrust at liftoff."
William responds:
Sure, Andy, the SRBs will lift the Shuttle and themselves off the
pad. I even carried out the calculation determining the gee forces at
lift-off and the gee-forces at burn-out.
That's not the issue. The issue is final speed.
An object must move at around 6.9 km/sec to maintain a stable orbit
around the Earth. Throw a ball - it follows an arc. Throw it faster
and the arc flattens out and the ball goes farther. Without air
resistance, you could throw a ball so that its arc has a radius of
curvature equal to the radius of the Earth - this is orbital speed;
Vo = sqrt(G*M/r)
Where G = gravitational constant
M = mass of the Earth
r = radius of orbit (must be greater than radius of
Earth)
Vo = orbital velocity
Like I said you need to go about 6.9 km/sec
http://en.wikipedia.org/wiki/Orbital_velocity
Now, the velocity of a rocket propelled projectile is given by the
Tsiolkovsky equation
Vf = Ve * LN(mi/mf)
Vf = final velocity
Ve = exhaust velocity
LN( ) = natural log function
mf = final mass
mi = initial mass
The Ve of the SRBs is around 2.6 km/sec
The mi/mf of the SRB/Shuttle configuration is 5.2
So, the Vf of the configuration is 2.6*LN(5.2) = 4.3 km/sec
This is LESS than orbital velocity. So, you cannot get into orbit
using three SRBs attached to a shuttle. The best you can do is lift
off, and arc over the Atlantic and land somewhere in the Indian Ocean.
> Ref: http://en.wikipedia.org/wiki/Space_Shuttle_main_engine
>
> 3 * 1.8 = 5.4 million newtons, less than half of the thrust of one SRB.
You erase the calculation I did. The SRBs can certainly lift the
Shuttle off the pad, they cannot however place it in orbit. They
cannot even place themselves in orbit. That's because the SRBs have
an exhaust velocity of 2.6 km/sec and they must achieve 6.9 km/sec
minimum - with air and gravity drag losses - 8.3 km/sec or so. It
cannot be done - which is WHY NASA used an ET to carry the higher
performing oxygen/hydrogen fuel which when burned in the SSME has an
exhaust velocity of 4.5 km/sec.
> The 786.5 tonne filled tank has a greater mass than the vehicle,
That's right. THAT'S THE WHOLE POINT!!!
The SRB when filled with solid rocket fuel has greater mass than the
vehicle too!
That's the point!
Propellant weights many times that of the payload are needed to attain
orbit.
Grok this and you are well on your way to understanding how rockets
work.
It is useful to compare and contrast two rocket types a solid rocket
with an exhaust speed of 2.6 km/sec and a liquid rocket with an
exhaust speed of 4.5 km/sec.
We'll fashion BOTH rockets which will be designed to attain a stable
orbital speed after air and gravity drag losses - both will therefore
achieve a final ideal speed of 9.2 km/sec - the same now achieved by
the Space Shuttle.
Alright, how much of the rocket must be propellant? This is the point
you missed Andy.
For the solid rocket the propellant fraction (u) is;
u = 1 - 1/EXP(Vf/Ve) = 1 - 1/EXP(9.2/2.6) = 0.97094
97.1% of the rocket must be propellant, leaving only 2.9% of the
rocket as structure and payload.
The Shuttle SRBs which are the best in the business have a propellant
fraction of 0.8539 - 85.4% - so they cannot attain orbit! Since they
cannot attain orbit lifting nothing - they cannot attain orbit lifting
something.
Gross Mass: 589,670 kg (1,299,990 lb).
Empty Mass: 86,183 kg (190,000 lb).
(589,670 - 86,183)/589,670 = 0.8539
This is why staging is used.
When we divide the rocket into three stages, we take the cube root of
the dimensionless propellant fraction - so, things get considerably
better.
Its best to go back to basics and figure out things by dividing the
velocity by the number of stages. So, we wanted to design a rocket
that achieved 9.2 km/sec therefore dividing that figure by 3 gets us
3.07 km/sec. This is the design speed for EACH of the 3 stages.
(we'll leave optimization aside since this takes calculus of
variations to solve it -lets get the basics first)
The SRB propellant has an exhaust velocity of 2.6 km/sec so the
propellant fraction needed to achieve this reduced speed is;
u = 1 - 1/exp(3.07/2.6) = 0.6929575
that's abut 69.3% propellant. Since 14.6% is structure, this leaves
1 - 0.693 - 0.146 = 0.161
16.1% of the original mass as payload!
So going back to the SRB we have 589,670 kg. The SRB lifts a SMALLER
mini-SRB atop its nose, and fires it after it burns out.
What's the size of it?
0.161 /(0.693 + 0.146) - 0.191895 ~ 0.191
It's 19.1% of the SRB weight or 112,626 kg.
589,670 * 0.191 = 112,626
This is the second stage third stage and payload.
19.1% of THAT weight is the third stage, and 19.1% of that payload is
the payload you can put on orbit with SRB technology;
705,207 kg - at lift off
589,670 kg - stage 1 (SRB adapted to first stage)
94,494 kg - stage 2 (Castor series adapted to 2nd stage)
18,132 kg - stage 3
2,911 kg - payload
So, this is how you can do it. And thrust isn't a problem as you
pointed out. Shaping the cavity so that burning surface is reduced as
the solid burns allows us to control thrust - so, you're wrong in
saying that you cannot control the thrust of a solid rocket. You can
- you program it in by controlling burn rates and shape of the burning
surface in the solid. This lets you maintain optimal thrust as you
fly your Goddard ascent curve.
Now, lets look at the liquid fuel technology represented by the ET and
SSME combination.
Exhaust velocity is 4.5 km/sec for this system and final velocity is
the same 9.2 km/sec. Applying the same rocket equation as we did
before using these performance figures obtains;
u = 1 - 1/exp(9.2/4.5) = 0.8705479 ~ 0.871
87.1% propellant fraction.
The external tank figures are;
Gross Mass: 750,975 kg (1,655,616 lb).
Empty Mass: 29,930 kg (65,980 lb).
So, u = (750,975 - 29,930) / 750,975 = 0.9602
around 96%
The SSME produces 6,834.303 kN (1,536,412 lbf) thrust. So, TWO SSME
can lift an External Tank with sufficient acceleration to be useful.
Three SSME can do an even better job.
So, this is interesting. You only need 87.1% propellant fraction and
structure fraction is 4% - this leaves;
1 - 0.871 - 0.04 = 0.089
or 8.9% of the lift off weight to orbit. This is
0.089/(0.871+0.04) = 0.097695 ~ 0.097
9.7% of the ET by itself - so,
750,975 * 0.097 = 72,884 kg
Taking away the 16,600 kg for the three SSME this leaves 56,244 kg
payload - which is 20x more than the 3stage SRB based system.
Now lets do the same calculation for a 3 stage system using SSME/ET
technology. We want to design three stages, with the ET/SSME first
stage, each stage to achieve 3.07 km/sec.
Going back and calculating propellant fraction;
1 - 1/exp(3.07/4.5) = 0.49450757 ~ 0.495
49.5% propellant fraction. So, the total vehicle weight is
Gross Mass: 750,975 kg (1,655,616 lb).
Empty Mass: 29,930 kg (65,980 lb).
(750,975-29,930)/0.495 = 1,456,656 kg
So, the three SSME can lift the whole 3 stage assembly off the pad at
1.435 gees - which is nearly optimal.
But what we're interested in is how does staging affect payload to
orbit for a vehicle this size?
Well, subtract off the fully loaded first stage to get the remaining
stages
1,456,656 - 750,975 = 705,681 kg
A single SSME drives this to 0.99 gees - which at 3.07 km/sec -
downrange, and flying horizontally - is sufficient to carry the stage
through its flight cycle.
Multiply by 50.5% to get the third stage plus payload
705,681 * 0.505 = 353,369 kg
Subtract the two figures to get second stage weight
705,681 - 353,369 = 349,312 kg
Multiply the third stage plus payload mass by 50.5% to get the payload
portion
356,369 kg * 0.505 = 179,996 kg
1.8 TIMES the payload of the Saturn V moon rocket!!!
And nearly 100x the payload lofted to orbit by an SRB based vehicle.
Subtract the two figures to get the third stage weight
356,369 - 179,966 = 176,403 kg
Okay so lets put this in a table side by side, after doubling up the
SRB - using two SRBs for the first stage, and this doubles the
payload;
Vehicle Type SRB ET/SSME
One Stage to Orbit infeasible 72,884 kg
Take off mass - 823,859 kg
Three Stage to Orbit 5,822 kg 179,996 kg
Take off mass 1,410,314 kg 1,456,656 kg
So, for comparable sized vehicles, 31x the mass to orbit could be put
up by an ET/SSME system when compared to the SRB system.
Andy continues;
>idiot
> Mooky.
Not at all. I computed take off gee forces and burn out gee forces -
which do not tell you what the burn out speed is! The rocket equation
gives you that.
You repeated the gee force calculation and erased the rocket equation
calculation that gives you final speed.
> If you want to argue that take it up with a clown of your
> own pathetic mentality.
Andy, you are the one who is limited in their thinking. How fast does
a rocket go? The rocket expels material out the back to push the
rocket forward. How fast and how much determines how fast the rocket
goes. Using Newton's laws and looking at the problem of an
accelerating rocket we have
F = dm/dt * Ve = m * dV/dt
Where F = thrust
dm/dt = mass flow rate of propellant
Ve = exhaust velocity
m = mass of the rocket
dV/dt = acceleration of the rocket
re-arranging terms we have
(dm/dt)/m = (dv/dt)/Ve
and integrating we have
ln(m(t)/m) = V(t)/Ve + C
Starting at lift off for the initial point, and ending at burn out at
the last point - we solve this equation to obtain
ln(m(empty)/m) = V(final) / Ve
rearrange to obtain the form
Vf = Ve * ln (mf/mi)
to calculate the speed of a stage
rearrange to obtain the form
(mi-mf)/mi = u = 1 - 1/exp(Vf/Ve)
and there you have it.
Rockets
SRB Ve=2.6 km/sec
SSME Ve=4.5 km/sec
Missions
Orbital Vf= 9.2 km/sec
Escape Vf=12.0 km/sec
> So... yes it would!
The SRBs certainly have enough thrust to lift the Shuttle as you
describe, but the configuration you have outlined would only achieve
4.6 km/sec by the time they burned out. So, this would not be enough
for orbit. It would be enough to loft the Shuttle on a sub-orbital
trajectory over the Atlantic, and by gliding - it would land somewhere
in the Indian ocean.
The only way an SRB technology will send something to orbit is to
divide the task into stages. Two SRBs would loft 5.6 metric tons into
orbit by adding two more stages to solid stages to them as described.
The ET/SSME technology will send something 78 metric tons into orbit
without staging. If the same staging system is applied to the ET/SSME
that payload to orbit increases to nearly 180 metric tons -
interestingly the three SSME are enough to do the job and the size of
the rocket at take off is nearly the same 1,400 metric tons for
each.
This is a good example of how important exhaust velocity and mass
ratio is to good rocket design. The ET is an incredible system - and
so is the SRB. The ET though is hands down winner for large launcher
capability. The SRB derived vehicle can be used to loft smaller
vehicles payloads reliably as well. Given the cost of solids versus
hydrogen/oxygen - not cheaply however - assuming fully reusable
systems.
> Your statement is a fuckin' lie
In what respect? I carried out the thrust calculations and gave you
the gee forces at lift off and burnout. That's not the issue. The
issue is the speed needed to maintain orbit and the speed the rocket
is capable of imparting to the shuttle. Thrust is one factor. The
other factor is the amount of time that thrust can be delivered. An
SRB can deliver 1 pound of thrust for 260 seconds using 1 pound of
propellant. An SSME/ET can deliver 1 pound of thrust for 455 seconds
using 1 pound of propellant. haha - are you getting it? - The rocket
equation lets you calculate how fast a rocket will go given the
propellant fraction and exhaust speed of the rocket.
Exhaust speed is related to specific impulse by;
Vf = g0 * Isp
g0 = 9.802 m/sec/sec in SI units
= 32.2 ft/sec/sec in english units
> designed to leave someone with your
> dumbfuck impression and your inability to perform simple arthmetic is
> laughable.
I did the simple arithmetic. The SRB will lift the Shuttle when
configured as you describe. It will not attain orbit, it will achieve
at most 4.6 km/sec which will take it as far as the Indian ocean when
launched from Florida in an Easterly direction.
Two SRB lofting two smaller SRB stages is capable of putting up 5
metric tons to LEO with a 1,400 ton lift off mass.
A single ET by itself with 3 SSME at its base, is capable of putting
up over 70 metric tons to LEO with a 750 ton lift off mass.
A single ET operating as a first stage, with 3 SSME lifting it,
pushing two smaller stages - the second stage with a single SSME, and
the third stage with another SSME or some smaller system - like RL10
or J2 - is capable of putting up nearly 180 tons to LEO with a 1,400
ton lift off mass.
> > Three steerable re-usable SRBs producing 12.5 million newtons, EACH,
> > of thrust would lift the shuttle to orbit without its 26.5 tonne tank,
William wrote;
> No it wouldn't.
William notes: (my explanation of the rocket equation is deleted)
Andy replies:
> ======================================
>
> "Each engine can generate almost 1.8 meganewtons (MN) or 400,000 lbf of
> thrust at liftoff."
William responds:
Sure, Andy, the SRBs will lift the Shuttle and themselves off the
pad. I even carried out the calculation determining the gee forces at
lift-off and the gee-forces at burn-out.
That's not the issue. The issue is final speed.
An object must move at around 6.9 km/sec to maintain a stable orbit
around the Earth. Throw a ball - it follows an arc. Throw it faster
and the arc flattens out and the ball goes farther. Without air
resistance, you could throw a ball so that its arc has a radius of
curvature equal to the radius of the Earth - this is orbital speed;
Vo = sqrt(G*M/r)
Where G = gravitational constant
M = mass of the Earth
r = radius of orbit (must be greater than radius of
Earth)
Vo = orbital velocity
Like I said you need to go about 6.9 km/sec
http://en.wikipedia.org/wiki/Orbital_velocity
Now, the velocity of a rocket propelled projectile is given by the
Tsiolkovsky equation
Vf = Ve * LN(mi/mf)
Vf = final velocity
Ve = exhaust velocity
LN( ) = natural log function
mf = final mass
mi = initial mass
The Ve of the SRBs is around 2.6 km/sec
The mi/mf of the SRB/Shuttle configuration is 5.2
So, the Vf of the configuration is 2.6*LN(5.2) = 4.3 km/sec
This is LESS than orbital velocity. So, you cannot get into orbit
using three SRBs attached to a shuttle. The best you can do is lift
off, and arc over the Atlantic and land somewhere in the Indian Ocean.
> Ref: http://en.wikipedia.org/wiki/Space_Shuttle_main_engine
>
> 3 * 1.8 = 5.4 million newtons, less than half of the thrust of one SRB.
You erase the calculation I did. The SRBs can certainly lift the
Shuttle off the pad, they cannot however place it in orbit. They
cannot even place themselves in orbit. That's because the SRBs have
an exhaust velocity of 2.6 km/sec and they must achieve 6.9 km/sec
minimum - with air and gravity drag losses - 8.3 km/sec or so. It
cannot be done - which is WHY NASA used an ET to carry the higher
performing oxygen/hydrogen fuel which when burned in the SSME has an
exhaust velocity of 4.5 km/sec.
> The 786.5 tonne filled tank has a greater mass than the vehicle,
That's right. THAT'S THE WHOLE POINT!!!
The SRB when filled with solid rocket fuel has greater mass than the
vehicle too!
That's the point!
Propellant weights many times that of the payload are needed to attain
orbit.
Grok this and you are well on your way to understanding how rockets
work.
It is useful to compare and contrast two rocket types a solid rocket
with an exhaust speed of 2.6 km/sec and a liquid rocket with an
exhaust speed of 4.5 km/sec.
We'll fashion BOTH rockets which will be designed to attain a stable
orbital speed after air and gravity drag losses - both will therefore
achieve a final ideal speed of 9.2 km/sec - the same now achieved by
the Space Shuttle.
Alright, how much of the rocket must be propellant? This is the point
you missed Andy.
For the solid rocket the propellant fraction (u) is;
u = 1 - 1/EXP(Vf/Ve) = 1 - 1/EXP(9.2/2.6) = 0.97094
97.1% of the rocket must be propellant, leaving only 2.9% of the
rocket as structure and payload.
The Shuttle SRBs which are the best in the business have a propellant
fraction of 0.8539 - 85.4% - so they cannot attain orbit! Since they
cannot attain orbit lifting nothing - they cannot attain orbit lifting
something.
Gross Mass: 589,670 kg (1,299,990 lb).
Empty Mass: 86,183 kg (190,000 lb).
(589,670 - 86,183)/589,670 = 0.8539
This is why staging is used.
When we divide the rocket into three stages, we take the cube root of
the dimensionless propellant fraction - so, things get considerably
better.
Its best to go back to basics and figure out things by dividing the
velocity by the number of stages. So, we wanted to design a rocket
that achieved 9.2 km/sec therefore dividing that figure by 3 gets us
3.07 km/sec. This is the design speed for EACH of the 3 stages.
(we'll leave optimization aside since this takes calculus of
variations to solve it -lets get the basics first)
The SRB propellant has an exhaust velocity of 2.6 km/sec so the
propellant fraction needed to achieve this reduced speed is;
u = 1 - 1/exp(3.07/2.6) = 0.6929575
that's abut 69.3% propellant. Since 14.6% is structure, this leaves
1 - 0.693 - 0.146 = 0.161
16.1% of the original mass as payload!
So going back to the SRB we have 589,670 kg. The SRB lifts a SMALLER
mini-SRB atop its nose, and fires it after it burns out.
What's the size of it?
0.161 /(0.693 + 0.146) - 0.191895 ~ 0.191
It's 19.1% of the SRB weight or 112,626 kg.
589,670 * 0.191 = 112,626
This is the second stage third stage and payload.
19.1% of THAT weight is the third stage, and 19.1% of that payload is
the payload you can put on orbit with SRB technology;
705,207 kg - at lift off
589,670 kg - stage 1 (SRB adapted to first stage)
94,494 kg - stage 2 (Castor series adapted to 2nd stage)
18,132 kg - stage 3
2,911 kg - payload
So, this is how you can do it. And thrust isn't a problem as you
pointed out. Shaping the cavity so that burning surface is reduced as
the solid burns allows us to control thrust - so, you're wrong in
saying that you cannot control the thrust of a solid rocket. You can
- you program it in by controlling burn rates and shape of the burning
surface in the solid. This lets you maintain optimal thrust as you
fly your Goddard ascent curve.
Now, lets look at the liquid fuel technology represented by the ET and
SSME combination.
Exhaust velocity is 4.5 km/sec for this system and final velocity is
the same 9.2 km/sec. Applying the same rocket equation as we did
before using these performance figures obtains;
u = 1 - 1/exp(9.2/4.5) = 0.8705479 ~ 0.871
87.1% propellant fraction.
The external tank figures are;
Gross Mass: 750,975 kg (1,655,616 lb).
Empty Mass: 29,930 kg (65,980 lb).
So, u = (750,975 - 29,930) / 750,975 = 0.9602
around 96%
The SSME produces 6,834.303 kN (1,536,412 lbf) thrust. So, TWO SSME
can lift an External Tank with sufficient acceleration to be useful.
Three SSME can do an even better job.
So, this is interesting. You only need 87.1% propellant fraction and
structure fraction is 4% - this leaves;
1 - 0.871 - 0.04 = 0.089
or 8.9% of the lift off weight to orbit. This is
0.089/(0.871+0.04) = 0.097695 ~ 0.097
9.7% of the ET by itself - so,
750,975 * 0.097 = 72,884 kg
Taking away the 16,600 kg for the three SSME this leaves 56,244 kg
payload - which is 20x more than the 3stage SRB based system.
Now lets do the same calculation for a 3 stage system using SSME/ET
technology. We want to design three stages, with the ET/SSME first
stage, each stage to achieve 3.07 km/sec.
Going back and calculating propellant fraction;
1 - 1/exp(3.07/4.5) = 0.49450757 ~ 0.495
49.5% propellant fraction. So, the total vehicle weight is
Gross Mass: 750,975 kg (1,655,616 lb).
Empty Mass: 29,930 kg (65,980 lb).
(750,975-29,930)/0.495 = 1,456,656 kg
So, the three SSME can lift the whole 3 stage assembly off the pad at
1.435 gees - which is nearly optimal.
But what we're interested in is how does staging affect payload to
orbit for a vehicle this size?
Well, subtract off the fully loaded first stage to get the remaining
stages
1,456,656 - 750,975 = 705,681 kg
A single SSME drives this to 0.99 gees - which at 3.07 km/sec -
downrange, and flying horizontally - is sufficient to carry the stage
through its flight cycle.
Multiply by 50.5% to get the third stage plus payload
705,681 * 0.505 = 353,369 kg
Subtract the two figures to get second stage weight
705,681 - 353,369 = 349,312 kg
Multiply the third stage plus payload mass by 50.5% to get the payload
portion
356,369 kg * 0.505 = 179,996 kg
1.8 TIMES the payload of the Saturn V moon rocket!!!
And nearly 100x the payload lofted to orbit by an SRB based vehicle.
Subtract the two figures to get the third stage weight
356,369 - 179,966 = 176,403 kg
Okay so lets put this in a table side by side, after doubling up the
SRB - using two SRBs for the first stage, and this doubles the
payload;
Vehicle Type SRB ET/SSME
One Stage to Orbit infeasible 72,884 kg
Take off mass - 823,859 kg
Three Stage to Orbit 5,822 kg 179,996 kg
Take off mass 1,410,314 kg 1,456,656 kg
So, for comparable sized vehicles, 31x the mass to orbit could be put
up by an ET/SSME system when compared to the SRB system.
Andy continues;
>idiot
> Mooky.
Not at all. I computed take off gee forces and burn out gee forces -
which do not tell you what the burn out speed is! The rocket equation
gives you that.
You repeated the gee force calculation and erased the rocket equation
calculation that gives you final speed.
> If you want to argue that take it up with a clown of your
> own pathetic mentality.
Andy, you are the one who is limited in their thinking. How fast does
a rocket go? The rocket expels material out the back to push the
rocket forward. How fast and how much determines how fast the rocket
goes. Using Newton's laws and looking at the problem of an
accelerating rocket we have
F = dm/dt * Ve = m * dV/dt
Where F = thrust
dm/dt = mass flow rate of propellant
Ve = exhaust velocity
m = mass of the rocket
dV/dt = acceleration of the rocket
re-arranging terms we have
(dm/dt)/m = (dv/dt)/Ve
and integrating we have
ln(m(t)/m) = V(t)/Ve + C
Starting at lift off for the initial point, and ending at burn out at
the last point - we solve this equation to obtain
ln(m(empty)/m) = V(final) / Ve
rearrange to obtain the form
Vf = Ve * ln (mf/mi)
to calculate the speed of a stage
rearrange to obtain the form
(mi-mf)/mi = u = 1 - 1/exp(Vf/Ve)
and there you have it.
Rockets
SRB Ve=2.6 km/sec
SSME Ve=4.5 km/sec
Missions
Orbital Vf= 9.2 km/sec
Escape Vf=12.0 km/sec
> So... yes it would!
The SRBs certainly have enough thrust to lift the Shuttle as you
describe, but the configuration you have outlined would only achieve
4.6 km/sec by the time they burned out. So, this would not be enough
for orbit. It would be enough to loft the Shuttle on a sub-orbital
trajectory over the Atlantic, and by gliding - it would land somewhere
in the Indian ocean.
The only way an SRB technology will send something to orbit is to
divide the task into stages. Two SRBs would loft 5.6 metric tons into
orbit by adding two more stages to solid stages to them as described.
The ET/SSME technology will send something 78 metric tons into orbit
without staging. If the same staging system is applied to the ET/SSME
that payload to orbit increases to nearly 180 metric tons -
interestingly the three SSME are enough to do the job and the size of
the rocket at take off is nearly the same 1,400 metric tons for
each.
This is a good example of how important exhaust velocity and mass
ratio is to good rocket design. The ET is an incredible system - and
so is the SRB. The ET though is hands down winner for large launcher
capability. The SRB derived vehicle can be used to loft smaller
vehicles payloads reliably as well. Given the cost of solids versus
hydrogen/oxygen - not cheaply however - assuming fully reusable
systems.
> Your statement is a fuckin' lie
In what respect? I carried out the thrust calculations and gave you
the gee forces at lift off and burnout. That's not the issue. The
issue is the speed needed to maintain orbit and the speed the rocket
is capable of imparting to the shuttle. Thrust is one factor. The
other factor is the amount of time that thrust can be delivered. An
SRB can deliver 1 pound of thrust for 260 seconds using 1 pound of
propellant. An SSME/ET can deliver 1 pound of thrust for 455 seconds
using 1 pound of propellant. haha - are you getting it? - The rocket
equation lets you calculate how fast a rocket will go given the
propellant fraction and exhaust speed of the rocket.
Exhaust speed is related to specific impulse by;
Vf = g0 * Isp
g0 = 9.802 m/sec/sec in SI units
= 32.2 ft/sec/sec in english units
> designed to leave someone with your
> dumbfuck impression and your inability to perform simple arthmetic is
> laughable.
I did the simple arithmetic. The SRB will lift the Shuttle when
configured as you describe. It will not attain orbit, it will achieve
at most 4.6 km/sec which will take it as far as the Indian ocean when
launched from Florida in an Easterly direction.
Two SRB lofting two smaller SRB stages is capable of putting up 5
metric tons to LEO with a 1,400 ton lift off mass.
A single ET by itself with 3 SSME at its base, is capable of putting
up over 70 metric tons to LEO with a 750 ton lift off mass.
A single ET operating as a first stage, with 3 SSME lifting it,
pushing two smaller stages - the second stage with a single SSME, and
the third stage with another SSME or some smaller system - like RL10
or J2 - is capable of putting up nearly 180 tons to LEO with a 1,400
ton lift off mass.