What Is A Likely Propellant/Fuel To Ship Ratio For An SSTO?

[Spacescifi]

It would seem... at least based on a previous conversation with sevenperforce, that a 10% fuel to ship ratio is not happening for an SSTO even if or especially if it can make orbit on it's own.... at least if it is a rocket.

I imagine an Orion could still manage with such ratios only because it is overpowered with thrust.

 

I suppose a 10% fuel to ship ratio is seen as some sort of scifi ideal.... but it seems the lower the fuel ratio the higher the energy you need to react with it to get to orbit... but engines can only take so much energy before destroying themselves. The alternative is burning a higher ratio of fuel to compensate for a less energetic reaction.... which is what we do with rockets in real life.... which thus favors staging over SSTOs.

In space a 10% fuel to ship ratio is possible, but again unless it is a very high energy reaction your mileage won't be very far.

 

Interestingly this leads to a conclusion I have made.... which you can feel free to correct me on about scifi SSTOs... provided they use some sort of advanced rocketry based on stuff we know like antimatter thermal rocketry.

My Conclusion: If you have a 10% fuel/propellant ratio and are using antimatter thermal rocketry for an SSTO you have a few options for dealing with waste heat.

1. Closed cycle cooling would require more of the ship to be engine and radiators than otherwise simply to deal with waste heat.

2. Don't. Just use antimatter project Orion propulsion AKA as chucking out photon torpedoes instead of nukes.

Open cycle rocket cooling is not an option since you would need more than 10% fuel ratios to reject all the waste heat from the antimatter thermal rocketry.

 

The lighter weight an SSTO is the easier it is to get to orbit powered by antimatter thermal rocketry... but ironically the waste heat is so high  that your payload would be relatively small since much of the ship would have to be engine and radiators for closed cycle cooling.

 

And that is the final nail in the coffin for an even somewhat realistic SSTO. Pointless as you cannot maximize payload to fuel ratio no matter what you do so long you hold to a 10% fuel/propellant ration.

All because of waste heat!

Why bother SSTOing if your payload will be really small compared to your engine/radiators.

May as well two-stage.

 

Exceptions would be airbreathing scramjet antimatter thermal rocket airplane SSTOs, but that has it's own problems just trying to avoid not burning up in the atmosphere.

 

Edited November 10, 2022 by Spacescifi

[kerbiloid]

fuel = 95%
cargo = 1%
random stuff = 4%

7 hours ago, Spacescifi said:

I imagine an Orion could still manage with such ratios only because it is overpowered with thrust.

Orion has the same TWR as the upper stage of any rocket.

Just to not fall down.

[Exoscientist]

 Dense propellant first stages already get well better than 10% structural fraction, i.e., 90% propellant fraction. For a long time it was felt a SSTO had to use a light fuel such as hydrogen because it had the highest Isp, ca. 450 s. But more careful analysis showed actually dense propellants would be better for a SSTO because a big component of the dry mass of a rocket is tankage and dense propellants, such as kerosene or methane, can carry more fuel for the same size tanks.( More on this below) 

 As an example of a first stage with very high propellant fraction, or said the other way, low structural fraction, take a look at the Falcon 9 first stage:

Type	Falcon 9 FT Stage 1
Length	42.6 m (47m w/ Interstage)
Diameter	3.66 m
Inert Mass	~22,200 kg (est.)
Propellant Mass	411,000 kg (According to FAA)
Fuel	Rocket Propellant 1
Oxidizer	Liquid Oxygen
LOX Mass	287,430 kg
RP-1 Mass	123,570 kg
LOX Volume	234,700 l
RP-1 Volume	143,900 l
LOX Tank	Monocoque
RP-1 Tank	Stringer & Ring Frame
Material	Aluminum-Lithium
Interstage Length	4.5 m (est.)
Guidance	From 2nd Stage
Tank Pressurization	Heated Helium
Propulsion	9 x Merlin 1D
Engine Arrangement	Octaweb

https://spaceflight101.com/spacerockets/falcon-9-ft/

  This is a structural fraction of 22,200/(22,200 + 411,000) = .05, 5%, so a propellant fraction of 95%. 

 However, the sea level Merlins don't have a very good Isp at vacuum ~311 s, so as an SSTO would get minimal payload to orbit , if any.

 Here's an analysis that shows a dense propellant SSTO can carry more payload to orbit than hydrogen fueled:

 

---

From: [email protected] (burnside)
Newsgroups: sci.space.policy
Subject: A LO2/kerosene SSTO rocket design, w/o AOL
Date: 2 Feb 1997 15:11:33 GMT

A LO2/Kerosene SSTO Rocket Design (long)

Mitchell Burnside Clapp
Pioneer Rocketplane

(view with a fixed pitch font such as courier or monaco)

Abstract

The NASA Access to Space LO2/hydrogen single stage to orbit
rocket was examined, and the configuration reaccomplished
with LO2/kerosene as the propellants. Four major changes
were made in assumptions. First, the aerodynamic
configuration was changed from a wing with winglets to a
swept wing with vertical tail. The delta-V for ascent
was as a result recalculated, yielding a lower value due to
different values for drag and gravity losses. The engines
were changed to LO2/kerosene burning NK-33 engines, which
have a much lower Isp than SSME-type engines used in the
access to space study, but also have a much higher
thrust-to-weight ratio. The orbital maneuvering system on
the Access to Space Vehicle was replaced with a pump-fed
system based on the D-58 engine used for that purpose now
on Proton stage 4 and Buran. Finally, the wing of the
vehicle was allowed to be wet with fuel, which is a
reasonable practice with kerosene but more controversial
with oxygen or hydrogen. Additionally, in order to reduce
the technology development needed, the unit weights of the
tankage were allowed to increase by 17 percent.

After the design was closed and all the weights
recalculated, the empty weight of the LO2/kerosene vehicle
was 35.6% lighter than its hydrogen fuelled counterpart.

Introduction

NASA completed a study in 1993 called Access to Space, the
purpose of which was to consider what sort of vehicle
should be operated to meet civil space needs in the future.
The study had three teams to evaluate three different broad
categories of options. The Option 3 team eventually settled
on a configuration called the SSTO/R. This vehicle was a
LO2/hydrogen vertical takeoff horizontal landing rocket.
The mission of the Access to Space vehicle was to place a
25,000 pound payload in a 220 n.mi. orbit inclined at 51.6
degrees. The vehicle had a gross liftoff weight of about
2.35 million pounds. The thrust at liftoff was 2.95 million
pounds, for a takeoff thrust to weight ratio of 1.2. The
empty weight of the vehicle was 222,582 pounds, and the
propellant mass fraction (defined here as
[GLOW-empty]/GLOW) was 90.5%.

Main power for this vehicle was provided by seven SSME
derivative engines, with the nozzle expansion ratio reduced
to 50. This resulted in an Isp reduction from 454 to 447.3
seconds. Each engine weighed 6,790 lbs, for an engine sea
level thrust to weight ratio of 62.

Aerodynamically the vehicle was fairly squat, with a
fineness ratio (length:diameter) of 5. The overall length
of the vehicle was 173 feet and its diameter was 34.6 feet.
It had a single main wing (dry of all propellants) of about
4,200 square feet total area, augmented by winglets for
directional control at reentry. The landing wing loading
was about 60 lb/ft2. The oxygen tank was in the nose
section. The payload was mounted transversely between the
oxygen and hydrogen tanks, and was 15 feet in diameter and
30 feet long.

This design exercise was among the most thorough ever
conducted of a single stage to orbit LO2/LH2 VTHL rocket.
It was probably the single greatest factor in convincing
the space agency that single stage to orbit flight was
feasible and practical, to borrow from the title of Ivan
Bekey's paper of the same name.

A LO2/kerosene alternative

A number of people have been asserting for some time that
higher propellant mass fractions available from dense
propellants may make single stage to orbit possible with
those propellants also. The historical examples of the
extraordinary mass fractions of the Titan II first stage,
the Atlas, and the Saturn first stage are all persuasive.
Further, denser propellants lead to higher engine thrust to
weight ratios, for perfectly understandable hydraulic
reasons.

It has not usually been observed that higher density also
leads to significant reductions in required delta-v.
There are two major reasons that this is so. First, the
reduction in volume leads to a smaller frontal area and
lower drag losses. The second, and more significant, reason
is that the gravity losses are also reduced. This is because
the mass of the vehicle declines more rapidly from its
initial value. The gravity losses are proportional to the
mass of the vehicle at any given time, and hence the
vehicle reaches its limit acceleration speed faster.

NASA itself has implicitly recognized this effect. When the
Access to Space Option 3 team examined tripropellant
vehicles, the delta-v to orbit derived from their work was
29,127 ft/sec, for precisely the reasons described in the
previous paragraph. This compares to a delta-v of 30,146
ft/s for the hydrogen-only baseline, as reported in a
briefing by David Anderson of NASA MSFC dated 6 October
1993. To be clear, these delta-v numbers include the back
pressure losses, so that no "trajectory averaged Isp"
number is used. They did not, however, report any results
for kerosene-only configurations.

To come to a more thorough understanding of the issues
involved in SSTO design, I have used the same methodology
as the Access to Space team to develop compatible numbers
for a LO2/kerosene SSTO. There are four major changes in
basic assumption between the two approaches, which I will
identify and justify here:

1: The ascent delta-v for the LO2/kerosene vehicle is
29,100 ft/sec, rather than 29,970 ft/sec. The reason for
this is argued above, but I ran POST to verify this value,
just to be sure. The target orbit is the same: 220 n.mi.
circular at 51.6 degrees inclination. The detailed weights
I have for the NASA vehicle are based on a delta-v of
29,970 ft/sec rather than the 30,146 ft/sec reported in
Anderson's work, but I prefer to use the values more
favourable to the hydrogen case to be conservative. The
optimum value of thrust to weight ratio turns out to be
slightly less than the hydrogen vehicle: 1.15 instead of
1.20.

2: The aerodynamic configuration is that of Boeing's RASV.
Without arguing whether this is optimal, the fineness ratio
of 8.27 and large wing lead to a much more airplane-like
layout, better glide and crossrange performance, and
reduced risk. The single vertical tail is simpler and safer
than winglets as well. Extensive analysis has justified the
reentry characterisitics of this aircraft. The wing is
assumed to be wet with the kerosene fuel, as is common on
most aircraft. The fuel is also present in the wing
carry-through box. The payload is carried over the wing
box, and the oxidizer tank is over the wing. This avoids
the need for an intertank, which in the NASA Access to
Space design is nearly 6,600 pounds.

3. The main propulsion system is the NK-33. The engine has
a sea level thrust of 339,416 lbs, a weight of 2,725 lbs
with gimbal, and a vacuum Isp of 331 seconds. Furthermore,
it requires a kerosene inlet pressure of only 2 psi
absolute, which dramatically reduces the pressure required
in the wing tank. It also operates with a LO2 pressure at
the inlet of only 32 psi. The comparable values for the
SSME are about 50 psi for both propellants. This will have
a substantial effect on the pressurization system weight.

4. The OMS weight is based on the D-58 engine. This engine
is used for the Buran OMS system and the Proton stage 4. As
heavy as it is the Isp is an impressive 354 seconds. NASA's
vehicle used a pressure fed OMS, which is a sensible design
choice if you're stuck with hydrogen and you wish to
minimize the number of fluids aboard the vehicle. But
because both oxygen and kerosene are space-storable, there
is no reason to burden the design with a heavy pressure fed
system.

Using the same methodology for calculating masses, and
accepting the subsystems masses as given in the Access to
Space vehicle, a redesign with oxygen and kerosene was
accomplished. The results appear in Table 1.

Table 1: Access to Space vehicle and LO2/kerosene
alternative

Name                           O2/H2      LO2/RP
Wing                          11,465      11,893 lb
Tail                           1,577       1,636 lb
Body                          64,748      33,741 lb
            Fuel tank         30,668           - lb
          Oxygen tank         13,273      17,271 lb
      Basic Structure         14,610      10,274 lb
  Secondary Structure          6,197       6,197 lb
Thermal Protection            31,098      21,238 lb
Undercarriage, aux. sys        7,548       5,097 lb
Propulsion, Main              63,634      36,426 lb
Propulsion, RCS                3,627       1,234 lb
Propulsion, OMS                2,280         823 lb
Prime Power                    2,339       2,339 lb
Power conversion & dist.       5,830       5,830 lb
Control Surface Actuation      1,549       1,549 lb
Avionics                       1,314       1,314 lb
Environmental Control          2,457       2,457 lb
Margin                        23,116      16,105 lb
Empty Weight                 222,582     141,682 lb

Payload                       25,000      25,000 lb

Residual Fluids                2,264       1,911 lb
     OMS and RCS               1,614       1,261 lb
      Subsystems                 650         650 lb
Reserves                       7,215       8,895 lb
     Ascent                    5,699       7,587 lb
        OMS                      679         541 lb
        RCS                      837         767 lb
Inflight losses               13,254      17,445 lb
         Ascent Residuals     10,984      15,175 lb
      Fuel Cell Reactants      1,612       1,612 lb
  Evaporator water supply        658         658 lb
Propellant, main           2,054,612   3,034,972 lb
     Fuel                    293,604     843,048 lb
     Oxygen                1,761,008   2,191,924 lb
Propellant, RCS                2,814       2,556 lb
     Orbital                   2,051       1,756 lb
     Entry                       763         800 lb
Propellant, OMS               19,357      15,452 lb
GLOW                       2,347,098   3,246,156 lb
Inserted Weight              292,486     211,185 lb
Pre-OMS weight               271,482     186,152 lb
Pre-entry Weight             252,125     170,700 lb
Landed Weight                251,362     169,900 lb
Empty weight                 222,582     141,682 lb

Sea Level Thrust           2,816,518   3,733,080 lb
Percent margin                 11.6%       12.8%
Assumed Isp(vac)               447.3       331.0 s
Ascent Delta-V                29,970      29,100 ft/s
OMS delta-V                    1,065         987 ft/s
RCS delta-V                      108         107 ft/s
Deorbit Delta-V                   44          53 ft/s
Reserves                       0.28%       0.25% lb/lb
Residuals                      0.53%       0.50% lb/lb
Wing Parameter                 4.56%       7.00% lb/lb
TPS parameter                 12.37%      12.50% lb/lb
Undercarriage parameter        3.00%       3.00% lb/lb
Wing Reference Area            4,189       5,528 ft2
Density of fuel                  4.4        50.5 lb/ft3
Density of oxygen               71.2        71.2 lb/ft3
Volume of fuel                66,276      16,694 ft3
Volume of oxygen              24,733      30,785 ft3
Fuel tank parameter             0.42           - lb/ft3
Oxygen tank parameter           0.48        0.56 lb/ft3

Some discussion of the results and justification is in
order.

The wing is about 40 percent heavier as a percentage of
landed weight than for the hydrogen fueled baseline. When
considered as a tank, it is about 60 percent heavier for
the volume of fuel it encloses. Its weight per exposed area
is about the same and the wing loading is half at landing.
No benefit is taken explicitly for the lack of a
requirement for kerosene tank cryogenic insulation.

The tail is assumed to have the same proportion of wing
weight for both cases. This is conservative for the
kerosene wehicle because its single vertical tail is
structurally more efficient.

The body of the kerosene vehicle has three components. The
oxidizer tank has an increased unit weight of about 17
percent. This is done in order to avoid the need for
aluminum-lithium, which was assumed in the Access to Space
vehicle. The basic structure group is unchanged, except
that the intertank is deleted and the thrust structure is
increased in proportion to the change in thrust level.
The secondary structure group is mostly payload support
related, and was not changed.

The thermal protection group is in both cases about 12.5%
of the entry weight. This works out to 1.107 lbs/ft2 of
wetted area for the kerosene vehicle, which is common to
many SSTO designs.

The undercarriage group is 3% of landed weight for both
vehicles. There is no benefit taken for reductions in gear
loads for the kerosene vehicle due to lower landing speed
and lower glide angle at landing.

The main propulsion group includes engines, base mounted
heat shield, and pressurization/feed weights. The engines
are far lighter for their thrust than SSME derivatives. The
pressurization weights are reduced in proportion to the
pressurized volume for the kerosene vehicle. No benefit is
taken for reduced tank pressure.

Here is as good a place as any to point out the erroneous
assertion that increased hydrostatic pressure is going to
lead to increased tankage weights. There is no requirement
for a particular ullage pressure except for the need to
keep the propellants liquid. It is the pressure at the base
of the fluid column rather than the top of the column that
is of engineering interest. The column of fluid exerts a
hydrostatic load on the base of the tank, but this load
does not typically exceed the much more adverse requirement
for engine inlet pressurization. For the kerosene vehicle,
the hydrostatic load at the base of the oxygen tank is 49
psi, which is compatible with the pressures normally seen
in oxygen tanks for rocket use. The load declines after
launch because the weight goes down faster than the
acceleration goes up.

The bottom line here is that dense propellants may require
you to alter a tank's pressurization schedule, but not to
overdesign the entire tank. Structures are sized by loads
and tankage for rockets is sized principally by volume, and if
the vehicle is small, by minimum gauge considerations.
This is not completely true for wet wings, however, as
discussed previously. In this particular example, there is
no need for high pressure in the wing tank either, because
of the low inlet pressure required by the NK-33.

The OMS group is the only other major change, as discussed
above. The reliable D-58 engine has been performing space
starts for decades and will serve well here. The
acceleration available from the OMS is about 0.12 g, which
is standard.

All the other weights are pushed straight across for the
most part. A brief inspection suggests that this is very
conservative. Control surface actuation requirements are
certainly less, electrical power requirements less, much
better fuel cells available than the phosporic acid type
assumed here, and reduced need for environmental control.
Nonetheless, rather than dispute any of these values it is
easier simply to accept them.

The margin is applied to all weight items at 15% execpt for
the engine group at 7.5%. The justification for this is that
the main and OMS engine weights are known to high accuracy.

The vehicle has an overall length of 1955 inches, and a
diameter of 236.4 inches. The wing has a leading edge sweep
of 55.5 degrees and a trailing edge sweep of -4.5 degrees.
Its reference area is 5,632 square feet, of which 3,992
square feet is exposed. The wing encloses 16,694 ft3 of
fuel, with a further 5% ullage. The carry-through is also
wet with fuel. The wing span is 1293 inches, and the taper
ratio is 0.13.

The payload bay has a maximum width and height of 15 feet.
It sits on top of the wing carry through box. The thrust
structure from the engines passes through and around the
payload bay to the forward LO2 tank. The payload bay is 30
feet in length. It has a pair of doors, the aft edge of
which is just forward of the vertical tail leading edge.

The engine section encloses 11 NK-33 engines, with a 4 - 3
- 4 layout. The engines are each 12.5 feet long, and
additional structure and subsystems take up another 6.5
feet.

The oxygen tank comprises the forward fuselage, which
encloses 30,785 ft3 of oxygen, with a further 5% ullage.
The length of the tank is about 100 feet. The ventral
surface of the tank is moderately flattened as it moves
aft, to fair smoothly with the wing lower surface. This
flattening reduces its length by about 5% with respect to a
strictly cylindrical layout. The aft edge of the oxygen tank
is about even with the forward payload bay bulkhead. A
compartment of about 13.9 feet provides room for some
subsystems and a potential cockpit in future versions.

Conclusion

The methods of the NASA Access to Space study were used to
design a single stage to orbit vehicle using existing
LO2/kerosene engines. An inspection of the final results
shows that the vehicle weighs about 36.5% less than its
hydrogen counterpart, with reductions in required
technology level and off the shelf engines. The center of
mass of the vehicle is about 61% of body length rather than
68% for the Access to Space vehicle, which should improve
control during reentry. The landing safety is considerably
improved by lower landing speed and better glide ratio.
Structural margins are greater overall. The vehicle
designed here appears to be superior in every respect:
smaller, lighter, lower required technology, improved
safety, and almost certainly lower development and
operations cost.

---

Edited November 11, 2022 by Exoscientist

[Exoscientist]

 

 By the way the two SSTO projects compared by Burnside would be a good Kerbal design project to see if they could really get that much payload to orbit.

 

  Bob Clark

Edited November 12, 2022 by Exoscientist

[Spacescifi]

3 hours ago, Exoscientist said:

 

 By the way the two SSTO projects compared by Burnside would be a good Kerbal design project to see if they could really get that much payload to orbit.

 

  Bob Clark

 

Well thank you Bob.

 

I was actually already aware that hydrogen required far bigger tanks due to being so low density.

Denser propellants basically mean you are hauling more energy per kllogram whereas with hydrogen you would need far more to even match it.

Ha... maybe someone will make doped molten lead propellant propellant someday.

 

I do not know if lead will react with anything for thrust, but I already know certain chemical reactions can be used rocketry with other metals.. such as alluminum.

[KSK]

This an odd thread in that the title is a question but the first post assumes an answer (10%) and then goes on to discuss some of the reasons why that answer is impractical or undesirable.

So why the insistence on 10%?

This is actually the kind of question that the (and stop me if you've heard this one before) rocket equation can answer very easily. You know the approximate delta-V required and antimatter solid core engines have a specific impulse of around 1000s.  Or, more properly, a couple of paper designs for antimatter engines that work in this scenario (sort of reasonable thrust to weight ratio with sufficient shielding to absorb the gamma rays and charged nasties created by antimatter annihilation), have an ISP of about 1000s.

Assume a delta-V of 10 km/s, an ISP of 1000s and a wet vehicle mass of 100 metric tons (because round numbers are easy). Plug that lot into the rocket equation and you get a dry vehicle mass of 36 tons.  So, from that, I conclude that a likely (for optimistic values of likely) propellant to ship ratio for an SSTO powered by an antimatter-thermal rocket is about 2:1, or 66% propellant.

Waste heat should not be a problem since the engine is a solid core engine and has been designed with the melting points of its various components in mind - hence the relatively modest ISP.

Lifting figures straight from Atomic Rockets, 11 tons of that dry vehicle mass is engine, and the engine has a thrust-to-weight ratio of 40:1.  Thrust to-weight ratio of the fully fueled vehicle is therefore a bit under 4:1, which should be more than enough oomph for requirements. 

In practice, if I was writing about such a beast, I would probably assume a wet vehicle mass of 200 tons, giving me 72 tons (or roughly 60 tons after the engine has been accounted for), of vehicle and payload to play around with. That would have a thrust-to-weight ratio of about 2:1 off the pad which is still fairly sporty. How much of those 60 tons would need to be set aside for antimatter storage, I have no idea.

Incidentally, if you do want an SSTO with 10% fuel, you need an engine with an ISP of about 10,000s. And an engine that can operate at that kind of efficiency with a high enough thrust-to-weight ratio to get off the ground -- is not going to happen outside of fiction. For pretty much the reasons you've already mentioned - waste heat.  Atomic Rockets has some hilarious figures for a gas-core antimatter engine. Ignoring the fact that sitting on top of a rocket powered by a ball of tungsten vapor which is being used as an antiproton target is probably a good working definition of insanity, the ISP for the beast is a mere 5000s (so nowhere near what we need for 10% fuel), and it weighs in at 182 tons, about half of which is radiators. Accordingly, it has a miserable thrust to weight ratio of 2.5 x10^-2, making it somewhat unsuitable for an SSTO.

 

Edited November 13, 2022 by KSK

[wumpus]

On 11/11/2022 at 1:51 AM, kerbiloid said:

fuel = 95%
cargo = 1%
random stuff = 4%

Orion has the same TWR as the upper stage of any rocket.

Just to not fall down.

Don't expect the 4% "random stuff" to be contain enough equipment to survive re-entry and landing on Earth.  For existing fuels, just use the rocket equation.  It will tell you that SSTO ain't gonna happen (at least re-usable SSTO).

But now with anti-matter (sufficiently "magic Isp" to allow SSTO).

Why worry about heat issues?  Assuming you are throwing the anti-matter into something more or less like a "combustion chamber" of a 21st century rocket, you should have exactly the same issues of lifting a similar massed rocket to orbit.  The important bit is to use open-loop cooling where the hot stuff is simply thrown out the end of the rocket (which is what  you want for rocket efficiency anyway).

PS: once you are in vacuum, the theoretical maximum for cooling in space should be ~60MW per radiators the size used on the ISS.  I computed this for a space combat thread and was shocked that it was a physical limit (going beyond this isn't just engineering, it would require a bunch of physical laws we know nothing about).

[Spacescifi]

46 minutes ago, wumpus said:

Don't expect the 4% "random stuff" to be contain enough equipment to survive re-entry and landing on Earth.  For existing fuels, just use the rocket equation.  It will tell you that SSTO ain't gonna happen (at least re-usable SSTO).

But now with anti-matter (sufficiently "magic Isp" to allow SSTO).

Why worry about heat issues?  Assuming you are throwing the anti-matter into something more or less like a "combustion chamber" of a 21st century rocket, you should have exactly the same issues of lifting a similar massed rocket to orbit.  The important bit is to use open-loop cooling where the hot stuff is simply thrown out the end of the rocket (which is what  you want for rocket efficiency anyway).

PS: once you are in vacuum, the theoretical maximum for cooling in space should be ~60MW per radiators the size used on the ISS.  I computed this for a space combat thread and was shocked that it was a physical limit (going beyond this isn't just engineering, it would require a bunch of physical laws we know nothing about).

 

This is why for all the improved efficiency of mini-mag orion... the original project Orion may still carry more payload, since the mini-mag would need a lot of radiator panels to approach the thrust of the original Orion... unless you are shipping less payload than the original to begin with.

Edited November 14, 2022 by Spacescifi

[kerbiloid]

1 hour ago, Spacescifi said:

 

This is why for all the improved efficiency of mini-mag orion... the original project Orion may still carry more payload, since the mini-mag would need a lot of radiator panels to approach the thrust of the original Orion... unless you are shipping less payload than the original to begin with.

Mini-Mag Orion still uses fission charges (by default made of curium, but plutonium is also acceptable).

Its reaction zone is still far behind the ship, so its heat problems are not much harder.

A fusion Mini-Mag is a theoretically possible further development.

[Spacescifi]

On 11/14/2022 at 11:41 AM, kerbiloid said:

Mini-Mag Orion still uses fission charges (by default made of curium, but plutonium is also acceptable).

Its reaction zone is still far behind the ship, so its heat problems are not much harder.

A fusion Mini-Mag is a theoretically possible further development.

 

Hmmm... are you implying that mini-mag Orion yeets the bomb out far behind the ship like with the original pusher plate before detonating it?

 

Perhaps I was unaware or wrong, but I presumed mini-mag orion used lasers on fuel pellets as they passed out the magnetic nozzle to create fusion blasts.

 

Or I assumed a bomb was blown up within the magnetic nozzle and the resulting plasma was propeled backward for thrust.

Nonetheless detonating a bomb farther out sounds reasonable because detobating it inside the nozzle would risk damaging it from debris.

Farther away detonations are safer for the magnetic nozzle I presume?

[kerbiloid]

3 hours ago, Spacescifi said:

Hmmm... are you implying that mini-mag Orion yeets the bomb out far behind the ship like with the original pusher plate before detonating it?

It's how it works. Just the "bombs" are primitive and remotely detonated by radiation pressure, they are far from the traditional nuke complexity.

3 hours ago, Spacescifi said:

Perhaps I was unaware or wrong, but I presumed mini-mag orion used lasers on fuel pellets as they passed out the magnetic nozzle to create fusion blasts.

It compresses the fission pellets to supercriticality with X-ray emistters.

[sevenperforce]

On 11/13/2022 at 1:06 PM, KSK said:

Accordingly, it has a miserable thrust to weight ratio of 2.5 x10^-2, making it somewhat unsuitable for an SSTO.

Only somewhat.

[KSK]

MiniMag Orion study report.

Wayback Machine (archive.org)

Uses magnetic compression of tiny fission pellets (see Figure 10) to initiate a fission reaction in those pellets.  Magnetic field is created by extremely high current pulses (as in 70 mega amp pulses). Plasma created by the exploding pellets is directed by a magnetic nozzle to generate thrust.

Pulsed power unit (essentially a large capacitor bank) is charged between pulses either by tapping the energy released by the pulses or by a steady state power unit (nuclear reactor?) which is used for engine start and to recharge the capacitor bank in the event of pellet misfires or non-fires.

The proposed thermal management system was non-trivial, weighing in at about 15.5 tons (see Tables I and II) but is a relatively low percentage of vehicle mass compared to the mass of the pulsed power unit (7 tons), steady state power unit (9 tons) and magnetic nozzle (103 tons) it.  So, waste heat is not the limiting factor on acceleration here. 

[Spacescifi]

6 hours ago, KSK said:

MiniMag Orion study report.

Wayback Machine (archive.org)

Uses magnetic compression of tiny fission pellets (see Figure 10) to initiate a fission reaction in those pellets.  Magnetic field is created by extremely high current pulses (as in 70 mega amp pulses). Plasma created by the exploding pellets is directed by a magnetic nozzle to generate thrust.

Pulsed power unit (essentially a large capacitor bank) is charged between pulses either by tapping the energy released by the pulses or by a steady state power unit (nuclear reactor?) which is used for engine start and to recharge the capacitor bank in the event of pellet misfires or non-fires.

The proposed thermal management system was non-trivial, weighing in at about 15.5 tons (see Tables I and II) but is a relatively low percentage of vehicle mass compared to the mass of the pulsed power unit (7 tons), steady state power unit (9 tons) and magnetic nozzle (103 tons) it.  So, waste heat is not the limiting factor on acceleration here. 

 

So for a manned mission with payload how does this compare to the original Orion?

What? Does mini-mag Orion out perform the original in every way?

 

Or is there anything the original project orion does better than mini-mag?

[sevenperforce]

2 hours ago, Spacescifi said:

Or is there anything the original project orion does better than mini-mag?

It gets off the ground.

[KSK]

4 hours ago, sevenperforce said:

It gets off the ground.

Yup. Although the thrust to weight ratio for mini-mag isn't bad. Mass at launch (including full propellant load) is 732 tons, thrust is 642 kN.  Pretty good for an engine with an ISP of 9,500 but 0.88 ms-² acceleration ain't enough to lift off from Earth.

[Edit, 0.88 ms-² not 0.88g as originally posted. Silly KSK - drink coffee first, then try basic arithmetic.]

@SpacescifiThe total mission delta-V for that particular version of mini-mag was 100km/s, with a payload of 100 tons. Even with such a high efficiency engine (and 9,500 is getting close to that 10,000 ISP required for an SSTO having 10% fuel that you mentioned earlier), the mini-mag vehicle was still about 2/3 propellant by mass. The rocket equation is unforgiving.

Addendum. If my coffee free maths is correct, that's an acceleration of 0.09g which can be sustained for a little over 31 hours.  Just to put the idea of a 3g constant acceleration drive into a bit of context.  

Edited November 18, 2022 by KSK

[Spacescifi]

29 minutes ago, KSK said:

Yup. Although the thrust to weight ratio for mini-mag isn't bad. Mass at launch (including full propellant load) is 732 tons, thrust is 642 kN.  Pretty good for an engine with an ISP of 9,500 but 0.88 ms-² acceleration ain't enough to lift off from Earth.

[Edit, 0.88 ms-² not 0.88g as originally posted. Silly KSK - drink coffee first, then try basic arithmetic.]

@SpacescifiThe total mission delta-V for that particular version of mini-mag was 100km/s, with a payload of 100 tons. Even with such a high efficiency engine (and 9,500 is getting close to that 10,000 ISP required for an SSTO having 10% fuel that you mentioned earlier), the mini-mag vehicle was still about 2/3 propellant by mass. The rocket equation is unforgiving.

Addendum. If my coffee free maths is correct, that's an acceleration of 0.09g which can be sustained for a little over 31 hours.  Just to put the idea of a 3g constant acceleration drive into a bit of context.  

 

It is too bad no one ever built mini-mag orion then.

 

As overall it is safer the the original.

 

I suppose launching it would be crazy expensive though so maybe that is why.

 

It would probably be constructed in orbit since launching it in one piece would probably require KSP level extreme boosters quantities.

 

 

So just to be clear, does mini-mag give greater mileage and an overall higher top speed than the original orion while the original trades high higher mileage for higher thrust at the cost of a lower top speed?

 

Because if that is the case then only in niche cases would you even want to have an original orion design built.

 

Seems to me that for go for heaviest payloads... like if you go crazy and make a mini-mag as heavy as a project orion battleship.... then a full comparison could be made.

 

So I presume the only difference is that mini-mag has a higher top speed but takes longer to get there, whereas the original orion has a lower top speed but gets there faster than the mini-mag?

 

Am I right or wrong?

[KSK]

Okay then.

Busting out my trusty copy of George Dyson's Project Orion (George is the son of Freeman Dyson who worked on Orion), the original design was intended to be capable of a total delta V of 20km/s (round trip to Moon levels of delta-V) with a mass ratio of 1.5:1. Interplanetary versions were designed with total delta-Vs of between 50 and 80 km/s with mass ratios of between 3 and 5.

The mass ratio is the mass of the fully fueled (for want of a better word) Orion divided by the mass of the empty Orion.  So, a mass ratio of 1.5 implies a vehicle that's about 30% propellant, whereas a mass ratio of 5 implies a vehicle that's about 80% propellant.

Compare that to Mini-mag Orion with its 100 km/s delta v and a mass ratio of about 3.3 (66% propellant).  That's considerably more efficient than Orion.

So yeah, you've got it about right. Assuming that you just point the ship at the nearest star and run the engine till you're out of propellant, you'll end up going faster in a Mini-Mag but it'll take you longer to get to that top speed.

Mileage doesn't really mean much in space travel - not as a means of comparing vehicle performance anyway. Simplistically, once you start moving at a given speed, you're not expending propellant to keep moving (like you are in a car), so you can travel as far as you like provided you're not worried about journey time. [Yes, yes, orbital mechanics and all that - I said 'simplistically' .]

The main reason you'd build an Orion over a Mini-Mag Orion is if you needed an actual launch vehicle. Orion will get you off the ground, Mini-Mag won't.  Also, as you said, you can build Orion on the ground, whereas Mini-Mag would probably need to be assembled on-orbit.

Also, you don't need to imagine a scaled up Mini-Mag to compare with Orion - at least not from a performance perspective. The rocket equation deals with ratios, so it doesn't matter whether you're talking about a 100 ton vehicle or a 100,000 ton vehicle - the amount of propellant needed as a fraction of total vehicle mass will just depend on the ISP of the engine. I suppose the payload fraction might be higher with the 100,000 ton vehicle but figuring that out would require knowing a whole lot more about the vehicle design, i.e. does a 100,000 ton Orion require a pusher plate / shock absorber assembly that's 1000 times heavier than that of a 100 ton Orion. My gut says 'probably not' which is why bigger Orion was almost always better, but then my gut knows both jack and squat about aerospace engineering.  

[sevenperforce]

I'm begging you, just learn the rocket equation.

You will quickly be able to deduce many important things all on your own.

Until then, please consult this handy guide.

Guide To Fictional Spaceships.

• For a spaceship that can jaunt back and forth between the surfaces of different planets in a matter of days:
  • You're going to need entirely new physics (hyperdrives, warp bubbles, etc.) at the very least.
  • Whatever new physics your spaceship uses, it is going to obviate any need for ordinary engines, propellant tanks, or other features commonly seen on rockets.
  • Your spaceship can be any shape you imagine and your engines can have any appearance you imagine, because you can make the rules for whatever new physics you want it to use.
  • Such trite and unimportant details like "waste heat" and "propellant capacity" and "delta-v" are necessary only if they are important to your plot, because your new physics obviates the Carnot cycle and the rocket equation.
• Do you want a spaceship that can travel from the orbit of one planet to the orbit of another planet in a matter of weeks, or travel back and forth between the orbits of various planets on a single propellant tank?
  • You're going to want a near-future high-energy low-thrust propulsion system like VASMIR or Mini-Mag or zeta-pinch, or a brute-force approach like orbital Orion or antimatter-thermal propulsion.
  • Your spaceship will be restricted to space alone and its engine will probably produce radiation you'll have to deal with.
  • Since you're stuck in space, radiators can handle waste heat, but they'll probably need to be quite large.
  • It's possible that your spaceship can be constructed in-space with a single conventional rocket launch, but multiple launches will probably be necessary.
  • Your propellant needs will be highly specialized, so you can't use ISRU, and your propellant will still probably make up half or more of your initial mass.
  • Your engine exhaust nozzle will not look like a conventional engine nozzle.
• Do you want a spaceship that can go from the surface of Earth to the surface of another world, or that can take significant payload to LEO and return intact, all in a single stage?
  • You're going to need a combination of chemical engines (for thrust) and nuclear thermal engines (for efficiency).
  • Propellant will be 90% or more of your liftoff mass.
  • If you're going to another world, the journey will use a Hohmann transfer, so it will take a long time.
  • If you're going beyond LEO, you're certainly not coming back home without refueling somewhere.
  • If your final destination has an atmosphere, you're going to have to figure out how to manage aerodynamic heating and maneuvering as well as descent and landing so that you can reach the ground intact.
  • You won't have to worry about waste heat because you'll be dumping all the heat into your exhaust.
  • Your exhaust nozzles will look reasonably normal.
• Do you want a spaceship that can fly around like an airplane in the atmosphere but can also be used as an orbital ferry?
  • You're going to want some combination of jet engines, rocket-combined-cycle engines, chemical rocket engines, and/or nuclear thermal engines.
  • Propellant will be at least 60-70% of your loaded weight, more if you're not using nuclear thermal propulsion.
  • It will need to a full propellant refill, either on the surface or in-flight, any time it wants to go to orbit.
  • You're not going to have margin for dual-axis thrust so it will either need to be a dedicated tailsitter or it will need to take off and land on a runway.
  • Your exhaust nozzles will look reasonably normal, and waste heat is not a problem.
• Do you have some other set of requirements outside of what is discussed above?
  • Learn the rocket equation and figure out what kind of performance you'll actually need, and go from there.

Hopefully that settles it.

[Spacescifi]

15 hours ago, KSK said:

Okay then.

Busting out my trusty copy of George Dyson's Project Orion (George is the son of Freeman Dyson who worked on Orion), the original design was intended to be capable of a total delta V of 20km/s (round trip to Moon levels of delta-V) with a mass ratio of 1.5:1. Interplanetary versions were designed with total delta-Vs of between 50 and 80 km/s with mass ratios of between 3 and 5.

The mass ratio is the mass of the fully fueled (for want of a better word) Orion divided by the mass of the empty Orion.  So, a mass ratio of 1.5 implies a vehicle that's about 30% propellant, whereas a mass ratio of 5 implies a vehicle that's about 80% propellant.

Compare that to Mini-mag Orion with its 100 km/s delta v and a mass ratio of about 3.3 (66% propellant).  That's considerably more efficient than Orion.

So yeah, you've got it about right. Assuming that you just point the ship at the nearest star and run the engine till you're out of propellant, you'll end up going faster in a Mini-Mag but it'll take you longer to get to that top speed.

Mileage doesn't really mean much in space travel - not as a means of comparing vehicle performance anyway. Simplistically, once you start moving at a given speed, you're not expending propellant to keep moving (like you are in a car), so you can travel as far as you like provided you're not worried about journey time. [Yes, yes, orbital mechanics and all that - I said 'simplistically' .]

The main reason you'd build an Orion over a Mini-Mag Orion is if you needed an actual launch vehicle. Orion will get you off the ground, Mini-Mag won't.  Also, as you said, you can build Orion on the ground, whereas Mini-Mag would probably need to be assembled on-orbit.

Also, you don't need to imagine a scaled up Mini-Mag to compare with Orion - at least not from a performance perspective. The rocket equation deals with ratios, so it doesn't matter whether you're talking about a 100 ton vehicle or a 100,000 ton vehicle - the amount of propellant needed as a fraction of total vehicle mass will just depend on the ISP of the engine. I suppose the payload fraction might be higher with the 100,000 ton vehicle but figuring that out would require knowing a whole lot more about the vehicle design, i.e. does a 100,000 ton Orion require a pusher plate / shock absorber assembly that's 1000 times heavier than that of a 100 ton Orion. My gut says 'probably not' which is why bigger Orion was almost always better, but then my gut knows both jack and squat about aerospace engineering.  

 

Thank you!

 

By 'mileage' I meant the ability of a spaceship to land upon or orbit closely around celestial bodies, instead of merely fly by them at high speed.

Once your propellant tank is empty you are done, both literally and figuratively in manned spaceflight.

Propellant is literally like the life blood of a spacecraft... only problem is that we are throwing it out to function as a spaceship.

So unless you wanna crash land or burn up on reentry, 'high mileage' is generally the better option for any manned flight.

3 hours ago, sevenperforce said:

I'm begging you, just learn the rocket equation.

You will quickly be able to deduce many important things all on your own.

Until then, please consult this handy guide.

Guide To Fictional Spaceships.

• For a spaceship that can jaunt back and forth between the surfaces of different planets in a matter of days:
  • You're going to need entirely new physics (hyperdrives, warp bubbles, etc.) at the very least.
  • Whatever new physics your spaceship uses, it is going to obviate any need for ordinary engines, propellant tanks, or other features commonly seen on rockets.
  • Your spaceship can be any shape you imagine and your engines can have any appearance you imagine, because you can make the rules for whatever new physics you want it to use.
  • Such trite and unimportant details like "waste heat" and "propellant capacity" and "delta-v" are necessary only if they are important to your plot, because your new physics obviates the Carnot cycle and the rocket equation.
• Do you want a spaceship that can travel from the orbit of one planet to the orbit of another planet in a matter of weeks, or travel back and forth between the orbits of various planets on a single propellant tank?
  • You're going to want a near-future high-energy low-thrust propulsion system like VASMIR or Mini-Mag or zeta-pinch, or a brute-force approach like orbital Orion or antimatter-thermal propulsion.
  • Your spaceship will be restricted to space alone and its engine will probably produce radiation you'll have to deal with.
  • Since you're stuck in space, radiators can handle waste heat, but they'll probably need to be quite large.
  • It's possible that your spaceship can be constructed in-space with a single conventional rocket launch, but multiple launches will probably be necessary.
  • Your propellant needs will be highly specialized, so you can't use ISRU, and your propellant will still probably make up half or more of your initial mass.
  • Your engine exhaust nozzle will not look like a conventional engine nozzle.
• Do you want a spaceship that can go from the surface of Earth to the surface of another world, or that can take significant payload to LEO and return intact, all in a single stage?
  • You're going to need a combination of chemical engines (for thrust) and nuclear thermal engines (for efficiency).
  • Propellant will be 90% or more of your liftoff mass.
  • If you're going to another world, the journey will use a Hohmann transfer, so it will take a long time.
  • If you're going beyond LEO, you're certainly not coming back home without refueling somewhere.
  • If your final destination has an atmosphere, you're going to have to figure out how to manage aerodynamic heating and maneuvering as well as descent and landing so that you can reach the ground intact.
  • You won't have to worry about waste heat because you'll be dumping all the heat into your exhaust.
  • Your exhaust nozzles will look reasonably normal.
• Do you want a spaceship that can fly around like an airplane in the atmosphere but can also be used as an orbital ferry?
  • You're going to want some combination of jet engines, rocket-combined-cycle engines, chemical rocket engines, and/or nuclear thermal engines.
  • Propellant will be at least 60-70% of your loaded weight, more if you're not using nuclear thermal propulsion.
  • It will need to a full propellant refill, either on the surface or in-flight, any time it wants to go to orbit.
  • You're not going to have margin for dual-axis thrust so it will either need to be a dedicated tailsitter or it will need to take off and land on a runway.
  • Your exhaust nozzles will look reasonably normal, and waste heat is not a problem.
• Do you have some other set of requirements outside of what is discussed above?
  • Learn the rocket equation and figure out what kind of performance you'll actually need, and go from there.

Hopefully that settles it.

Thanks... one day I may indeed learn the rocket equation... but current demands in my life require priorities associated with my own future more than anything else.

 

And yes I had come to realize scifi SSTOs like in the movies would require either a revolution in physics or just making it all up.... so I will make it up.

The main limits being a cap on max acceleration, since if you increase power to scifi engines to accelerate higher than the cap the engines will explode... thus the acceleration cap.

A 3g acceleration cap is useful for most things... though I think some have told me that places with gravity of 3g or higher the ship would struggle to escape from or not even be able to... at least not without disposable rocket boosters that do not have a 3g cap on their acceleration.

So no deep diving into Jupiter.... that is insane anyway lol.

Edited November 18, 2022 by Spacescifi

[KSK]

7 hours ago, Spacescifi said:

Thank you!

By 'mileage' I meant the ability of a spaceship to land upon or orbit closely around celestial bodies, instead of merely fly by them at high speed.

Once your propellant tank is empty you are done, both literally and figuratively in manned spaceflight.

Propellant is literally like the life blood of a spacecraft... only problem is that we are throwing it out to function as a spaceship.

So unless you wanna crash land or burn up on reentry, 'high mileage' is generally the better option for any manned flight.

Okay, that's cool - thanks for the explanation!

At the risk of being a terrible old pedant though, it may be worth starting to think of 'mileage' as a series of velocity changes (if you're not doing so already).  For example, an Apollo style trip to the Moon required one burn (velocity change) to break out of low Earth orbit, another to enter Lunar orbit, and a third to depart from Lunar orbit and head home. The landings of course required additional velocity changes.  In principle, one could make a fourth burn to re-enter Earth orbit and complete the round trip, but Apollo didn't have that capability.

A couple of advantages to the above in my opinion. 

Firstly - it's how everyone else thinks about spacecraft trajectories (or at least, it's better aligned to how everyone else thinks about spacecraft trajectories) so if you're looking stuff up about spaceflight or discussing spaceflight with other people, it's easier to be thinking in terms of 'series of velocity changes' rather than 'mileage' and having to mentally translate between the two all the time. Secondly, if you're writing space science fiction (unless you're going full on fictional physics as per @sevenperforce's post) then talking about velocity changes will sound more authentic than talking about mileage.

Personal anecdote. Back in school when I started learning about equations of motion in Physics, it made zero sense to me that distance, or displacement, was denoted by 's' rather than 'd' and all my early notes used 'd' instead. Eventually though, it just got to be too much of a pain to use 'd' when my textbooks and teacher were using 's', so I gave up and went with the standard 's' instead.