R
RGClark
Guest
Recently announced:
---------------------------------------------------------------------
Boeing proposes SSTO system for AF RBS program.
The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they
are proposing a 3-4 year technology readiness assessment that would
lead up to a demonstration of a X-37B type of system but would be
smaller. Wind tunnel tests have been completed. Davis says the system
would be a single stage capable of reaching low Earth orbit and, with
a booster, higher orbits. The system would return to Earth as a
glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers.
http://www.hobbyspace.com/nucleus/index.php?itemid=30110
---------------------------------------------------------------------
It is my contention that the reason why launch costs are so high, the
reason why we don't have passenger access to space as routine as say
trans-Pacific flights is that the idea has been promulgated that SSTO
is impossible. That is not the case. In fact it is easy, IF you do it
in the right way. The right way is summarized in one simple
statement:
If you use both weight optimized structures and highest efficiency
engines at the same time, then what you wind up with will be SSTO
capable whether you intend it to or not.
We all know that to get a good payload to space you want a high
efficiency engine. And we all know we want to use lightweight
structures so the weight savings can go to increased payload. So you
would think it would be obvious to use both these ideas to maximize
the payload to orbit, right?
And indeed both have been used together - for upper stages. Yet this
fundamentally obvious concept still has not been used for *first
stages*. It is my thesis that if you do this, then what you wind up
with will automatically be SSTO capable. This is true for either
kerosene fueled or hydrogen fueled stages.
Part of the misinformation that has been promulgated is that the mass
ratio for SSTO's is some impossible number. This is false. We've had
rocket stages with the required mass ratio's since the 60's, nearly 50
years, both for kerosene and hydrogen fueled. Another part of the
misinformation is that it would require some unknown high energy fuel
and engine to accomplish. This is false. The required engines have
existed since the 70's, nearly 40 years, both for kerosene and
hydrogen fueled.
What has NOT been done is to marry the two concepts together for
first stages. All you need to do is swap out the low efficiency
engines that have been used for the high mass ratio stages and replace
them with the high efficiency engines. It really is that simple.
This makes possible small, low cost orbital vehicles that could
transport the same number of passengers as the space shuttle, about 7,
but would have a comparable cost to a mid-sized business jet, a few
tens of millions of dollars.
Then once you have the SSTO's they make your staged vehicles even
better because you can carry greater payload when they are used for
the individual stages of the multi-staged vehicle.
In disseminating the false dogma that SSTO's are not possible it is
sometimes said instead that they are not practical because the payload
fraction is so small. Even this is false. And indeed this is just as
damaging as making the false statement they are not possible because
the statements are often conflated into meaning the same thing. So
when those in the industry make the statement they are not
"practical", meaning actually they are doable but not economical, this
becomes interpreted among many space enthusiasts and even many in the
industry as meaning it would require some revolutionary advance to
make them possible.
The fact that you can carry significant payload to orbit using SSTO's
can be easily confirmed by anyone familiar with the rocket equation.
To get a SSTO with significant payload using efficient kerosene
engines you need a mass ratio of about 20 to 1. And to get a SSTO with
significant payload using efficient hydrogen engines you need a mass
ratio of about 10 to 1. Both of these the high mass ratio stages and
the high efficiency engines for both kerosene and hydrogen have
existed for decades now.
See this list of rocket stages:
Stages Index.
http://www.friends-partners.org/partners/mwade/stages/staindex.htm
Among the kerosene-fueled stages you see that several among the Atlas
and Delta family have the required mass ratio. However, for the early
Atlas stages you have to be aware of the type of staging system they
used. They had drop-off booster engines and a main central engine,
called the sustainer that continued all the way to orbit. But even
when you take this into account you see these highly weight optimized
stages had surprisingly high mass ratios.
See for instance the Atlas Agena SLV-3:
SLV-3 Atlas / Agena B.
Family: Atlas. Country: USA. Status: Hardware. Department of Defence Designation: SLV-3.
Standardized Atlas booster with Agena B upper stage.
Specifications
Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg inclination trajectory.
Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass: 3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec. Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m. Propellants: Lox/Kerosene No Engines: 2. LR-89-5
Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg. Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length: 20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5
Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867 kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0 sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric acid/UDMH No Engines: 1. Bell 8081
http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm
Looking at only the gross mass/empty mass of stage 1, you would think this
stage had a mass ratio close to 50 to 1. But that is only including the
sustainer engine. The more relevant ratio would be when you add in the
mass of the jettisonable booster engines to the dry mass since they are
required to lift the vehicle off the pad. These are contained within the stage 0
mass at 3,174 kg. This makes the loaded mass now 117,026 + 3,174 =
120,200 and the dry mass 2,326 + 3,174 = 5,500 kg, for a mass ratio of
21.85.
But this was using the low efficiency engines available in the early
60's. Let's swap these out for the high efficiency NK-33 [1]. The
sustainer engine used was the LR-105-5 [2] at 460 kg. At 1,220 kg the
NK-33 weighs 760 kg more. So removing both the sustainer and booster
engines to be replaced by the NK-33 our loaded mass becomes 117,786 kg
and the dry mass 3,086 kg, and the mass ratio 38.2 (!).
For the trajectory-averaged Isp, notice this is not just the midpoint
between the sea level and vacuum value, since most of the flight to
orbit is at high altitude at near vacuum conditions. A problem with
doing these payload to orbit estimates is the lack of a simple method
for getting the average Isp over the flight for an engine, which
inhibits people from doing the calculations to realize SSTO is
possible and really isn't that hard. I'll use a guesstimate Ed Kyle
uses, who is a frequent contributor to NasaSpaceFlight.com and the
operator of the Spacelaunchreport.com site. Kyle takes the average Isp
as lying 2/3rds of the way up from the sea level value to the vacuum
value [3]. The sea level value of the Isp for the NK-33 is 297 s, and
the vacuum value 331 s. Then from this guesstimate the average Isp is
297 + (2/3)(331 - 297) = 319.667, which I'll round to 320 s.
Using this average Isp and a 8,900 m/s delta-V for a flight to orbit,
we can lift 4,200 kg to orbit:
320*9.8ln((117,786+4,000)/(3,086+4,000)) = 8,919 m/s. This is a
payload fraction of 3.3%, comparable to that of many multi-stage
rockets.
Note in fact that this has a very good value for a ratio that I
believe should be regarded as a better measure, i.e., figure of merit,
for the efficiency of a orbital vehicle. This is the ratio of the
payload to the total dry mass of the vehicle. The reason why this is a
good measure is because actually the cost of the propellant is a minor
component for the cost of an orbital rocket. The cost is more
accurately tracked by the dry mass and the vehicle complexity. Note
that SSTO's in not having the complexity of staging are also good on
the complexity scale.
For the ratio of the payload to dry mass you see this is greater than
1 for this SSTO. This is important because for every orbital vehicle I
looked at, and possibly for every one that has existed, this ratio is
going in the other direction: the vehicle dry mass is greater than the
payload carried. Often it is much greater. For instance for the space
shuttle system, the vehicle dry mass is more than 12 times that of the
payload.
This good payload fraction and even better payload to dry mass ratio
was just by using the engine in its standard configuration, no
altitude compensation. However, for a SSTO you definitely would want
to use altitude compensation. Dr. Bruce Dunn in his report "Alternate
Propellants for SSTO Launchers" [4] estimates an average Isp of 338.3
s for high performance kerosene engines when using altitude
compensation. Then we could lift 5,300 kg to orbit:
338.3*9.8ln((117,786+5,300)/(3,086+5,300)) = 8,906 m/s.
But kerosene is not the most energetic hydrocarbon fuel you could
use. Dunn in his report estimates an average Isp of 352 s for
methylacetyene using altitude compensation. This would allow a payload
of 6,300 kg : 352*9.8ln((117,526+6,300)/(3,086+6,300)) = 8,900 m/s.
Bob Clark
REFERENCES.
1.)NK-33.
http://www.friends-partners.org/partners/mwade/engines/nk33.htm
2.)LR-105-5.
http://www.friends-partners.org/partners/mwade/engines/lr1055.htm
3.)EELV Solutions for VSE.
Reply #269 on: 11/05/2007 09:20 PM
http://forum.nasaspaceflight.com/index.php?topic=10497.msg208875#msg208875
4.)Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm
Disclaimer: the citing of a particular reference should not be
construed as an endorsement by the cited authors of the viewpoint
expressed herein.
---------------------------------------------------------------------
Boeing proposes SSTO system for AF RBS program.
The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they
are proposing a 3-4 year technology readiness assessment that would
lead up to a demonstration of a X-37B type of system but would be
smaller. Wind tunnel tests have been completed. Davis says the system
would be a single stage capable of reaching low Earth orbit and, with
a booster, higher orbits. The system would return to Earth as a
glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers.
http://www.hobbyspace.com/nucleus/index.php?itemid=30110
---------------------------------------------------------------------
It is my contention that the reason why launch costs are so high, the
reason why we don't have passenger access to space as routine as say
trans-Pacific flights is that the idea has been promulgated that SSTO
is impossible. That is not the case. In fact it is easy, IF you do it
in the right way. The right way is summarized in one simple
statement:
If you use both weight optimized structures and highest efficiency
engines at the same time, then what you wind up with will be SSTO
capable whether you intend it to or not.
We all know that to get a good payload to space you want a high
efficiency engine. And we all know we want to use lightweight
structures so the weight savings can go to increased payload. So you
would think it would be obvious to use both these ideas to maximize
the payload to orbit, right?
And indeed both have been used together - for upper stages. Yet this
fundamentally obvious concept still has not been used for *first
stages*. It is my thesis that if you do this, then what you wind up
with will automatically be SSTO capable. This is true for either
kerosene fueled or hydrogen fueled stages.
Part of the misinformation that has been promulgated is that the mass
ratio for SSTO's is some impossible number. This is false. We've had
rocket stages with the required mass ratio's since the 60's, nearly 50
years, both for kerosene and hydrogen fueled. Another part of the
misinformation is that it would require some unknown high energy fuel
and engine to accomplish. This is false. The required engines have
existed since the 70's, nearly 40 years, both for kerosene and
hydrogen fueled.
What has NOT been done is to marry the two concepts together for
first stages. All you need to do is swap out the low efficiency
engines that have been used for the high mass ratio stages and replace
them with the high efficiency engines. It really is that simple.
This makes possible small, low cost orbital vehicles that could
transport the same number of passengers as the space shuttle, about 7,
but would have a comparable cost to a mid-sized business jet, a few
tens of millions of dollars.
Then once you have the SSTO's they make your staged vehicles even
better because you can carry greater payload when they are used for
the individual stages of the multi-staged vehicle.
In disseminating the false dogma that SSTO's are not possible it is
sometimes said instead that they are not practical because the payload
fraction is so small. Even this is false. And indeed this is just as
damaging as making the false statement they are not possible because
the statements are often conflated into meaning the same thing. So
when those in the industry make the statement they are not
"practical", meaning actually they are doable but not economical, this
becomes interpreted among many space enthusiasts and even many in the
industry as meaning it would require some revolutionary advance to
make them possible.
The fact that you can carry significant payload to orbit using SSTO's
can be easily confirmed by anyone familiar with the rocket equation.
To get a SSTO with significant payload using efficient kerosene
engines you need a mass ratio of about 20 to 1. And to get a SSTO with
significant payload using efficient hydrogen engines you need a mass
ratio of about 10 to 1. Both of these the high mass ratio stages and
the high efficiency engines for both kerosene and hydrogen have
existed for decades now.
See this list of rocket stages:
Stages Index.
http://www.friends-partners.org/partners/mwade/stages/staindex.htm
Among the kerosene-fueled stages you see that several among the Atlas
and Delta family have the required mass ratio. However, for the early
Atlas stages you have to be aware of the type of staging system they
used. They had drop-off booster engines and a main central engine,
called the sustainer that continued all the way to orbit. But even
when you take this into account you see these highly weight optimized
stages had surprisingly high mass ratios.
See for instance the Atlas Agena SLV-3:
SLV-3 Atlas / Agena B.
Family: Atlas. Country: USA. Status: Hardware. Department of Defence Designation: SLV-3.
Standardized Atlas booster with Agena B upper stage.
Specifications
Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg inclination trajectory.
Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass: 3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec. Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m. Propellants: Lox/Kerosene No Engines: 2. LR-89-5
Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg. Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length: 20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5
Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867 kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0 sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric acid/UDMH No Engines: 1. Bell 8081
http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm
Looking at only the gross mass/empty mass of stage 1, you would think this
stage had a mass ratio close to 50 to 1. But that is only including the
sustainer engine. The more relevant ratio would be when you add in the
mass of the jettisonable booster engines to the dry mass since they are
required to lift the vehicle off the pad. These are contained within the stage 0
mass at 3,174 kg. This makes the loaded mass now 117,026 + 3,174 =
120,200 and the dry mass 2,326 + 3,174 = 5,500 kg, for a mass ratio of
21.85.
But this was using the low efficiency engines available in the early
60's. Let's swap these out for the high efficiency NK-33 [1]. The
sustainer engine used was the LR-105-5 [2] at 460 kg. At 1,220 kg the
NK-33 weighs 760 kg more. So removing both the sustainer and booster
engines to be replaced by the NK-33 our loaded mass becomes 117,786 kg
and the dry mass 3,086 kg, and the mass ratio 38.2 (!).
For the trajectory-averaged Isp, notice this is not just the midpoint
between the sea level and vacuum value, since most of the flight to
orbit is at high altitude at near vacuum conditions. A problem with
doing these payload to orbit estimates is the lack of a simple method
for getting the average Isp over the flight for an engine, which
inhibits people from doing the calculations to realize SSTO is
possible and really isn't that hard. I'll use a guesstimate Ed Kyle
uses, who is a frequent contributor to NasaSpaceFlight.com and the
operator of the Spacelaunchreport.com site. Kyle takes the average Isp
as lying 2/3rds of the way up from the sea level value to the vacuum
value [3]. The sea level value of the Isp for the NK-33 is 297 s, and
the vacuum value 331 s. Then from this guesstimate the average Isp is
297 + (2/3)(331 - 297) = 319.667, which I'll round to 320 s.
Using this average Isp and a 8,900 m/s delta-V for a flight to orbit,
we can lift 4,200 kg to orbit:
320*9.8ln((117,786+4,000)/(3,086+4,000)) = 8,919 m/s. This is a
payload fraction of 3.3%, comparable to that of many multi-stage
rockets.
Note in fact that this has a very good value for a ratio that I
believe should be regarded as a better measure, i.e., figure of merit,
for the efficiency of a orbital vehicle. This is the ratio of the
payload to the total dry mass of the vehicle. The reason why this is a
good measure is because actually the cost of the propellant is a minor
component for the cost of an orbital rocket. The cost is more
accurately tracked by the dry mass and the vehicle complexity. Note
that SSTO's in not having the complexity of staging are also good on
the complexity scale.
For the ratio of the payload to dry mass you see this is greater than
1 for this SSTO. This is important because for every orbital vehicle I
looked at, and possibly for every one that has existed, this ratio is
going in the other direction: the vehicle dry mass is greater than the
payload carried. Often it is much greater. For instance for the space
shuttle system, the vehicle dry mass is more than 12 times that of the
payload.
This good payload fraction and even better payload to dry mass ratio
was just by using the engine in its standard configuration, no
altitude compensation. However, for a SSTO you definitely would want
to use altitude compensation. Dr. Bruce Dunn in his report "Alternate
Propellants for SSTO Launchers" [4] estimates an average Isp of 338.3
s for high performance kerosene engines when using altitude
compensation. Then we could lift 5,300 kg to orbit:
338.3*9.8ln((117,786+5,300)/(3,086+5,300)) = 8,906 m/s.
But kerosene is not the most energetic hydrocarbon fuel you could
use. Dunn in his report estimates an average Isp of 352 s for
methylacetyene using altitude compensation. This would allow a payload
of 6,300 kg : 352*9.8ln((117,526+6,300)/(3,086+6,300)) = 8,900 m/s.
Bob Clark
REFERENCES.
1.)NK-33.
http://www.friends-partners.org/partners/mwade/engines/nk33.htm
2.)LR-105-5.
http://www.friends-partners.org/partners/mwade/engines/lr1055.htm
3.)EELV Solutions for VSE.
Reply #269 on: 11/05/2007 09:20 PM
http://forum.nasaspaceflight.com/index.php?topic=10497.msg208875#msg208875
4.)Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm
Disclaimer: the citing of a particular reference should not be
construed as an endorsement by the cited authors of the viewpoint
expressed herein.
