Ok, here are a few that jumped out at me.
This is a baseless and most likely incorrect assumption. Even if your aerospike made out of wonder-materials performs as advertised you do not appear to have accounted for the structural modifications that would need to be made to the stage itself to accomidate the spike and repositioned engines.
I'll repeat what Urwumpe said...
*Which modern materials?
*Is it really only because of the materials or are different structural solutions also required?
*How probable?
*Where are you getting "half"?
You do realize that altitude compensation does not increase ISP correct? It simply cuts your ambient pressure losses. Have you actually plotted out your thrust curve from sea-level up through MAX-Q and back?
Speaking as an engineering student, I stopped and laughed at this point. :lol: The modifications you propose would require signifigant structural changes and re-working of the engine's auxillary components. (Fuel-pumps, cooling, etc...) Depending on how much re-design is actually required it might actually be easier to start from scratch.
Mass has very little to do with cost. The chief driver of $ per Kilogram is the number of man-hours required per launch. (and in the case of a re-usable booster recovery/refurbishment as well)
1.)Lightweight aerospikes.
Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
Part of the Integrated High Payoff
Rocket Propulsion Technology (IHPRPT) program, the efforts
of the directorate include:
1.Developing continuous fiber reinforced ceramic matrix
composites (CMCs) for actively cooled thrust chambers and
nozzles
2.Demonstrating the feasibility of a transpiration-cooled thrust
chamber
3. Evaluating ceramic matrix composites for radiation cooled
nozzles.
The goal is to develop and demonstrate these new technologies
so that they may be incorporated into future rocket
engines. Using lightweight ceramics has the potential to reduce
the weight of the combustion devices by up to 50%.
...
Table 1 lists the materials and type of construction of
numerous combustion devices, both historical and current. As
the table shows, the materials of choice (for all the engine manufacturers)
for combustion devices in large liquid fueled rocket
engines have historically been stainless steels, nickel-based
superalloys, and copper alloys. These materials are selected for
their high strength and high thermal conductivity in order
to cope with the stresses and extreme thermal environments
of rocket engines. Since these alloys also have high densities
(8-9 g/cm3), widespread reliance on them has traditionally
resulted in heavy engines.
Designers would like to reduce the weight of rocket
engines. A key performance criterion for engines is thrust-to-weight
ratio. Lighter engines and launch vehicles would allow
heavier payloads to be placed into orbit at a lower cost. One
path to lighter weight engines is replacement of conventional
high-density engine alloys with lightweight, high specific
strength ceramic composites. Two attractive candidates for
this application are carbon fiber reinforced silicon carbide
(C/SiC) and silicon carbide fiber reinforced silicon carbide
(SiC/SiC). These materials have low densities (2.0-2.4 g/cm3)
and high strengths that they maintain to relatively high temperatures
(2400-3000°F).
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf
Reportedly such lightweight, high temperature ceramics have already been used on hypersonic, scramjet test vehicles. Keep in mind also that for this application while you are adding on the weight of the aerospike you are also removing much of the weight of the nozzles.
For the size of the engines fitting within the Falcon 9 diameter around a central spike, remember most of the nozzles will be removed. As it is now, 8 engines are in a square pattern with the ninth engine in the center. So there are 4 corner engines with 1 engine between any pair of these corner engines. Then these engines between the corner engines can be moved outside instead to the circumscribed circle around the square. Taking into account the extra space without much of the nozzle width the ninth engines should fit in this space as well.
Furthermore, I realized that you don't need all 9 engines now. With the upper stage removed and with the smaller payload, and with the higher thrust and lighter weight of the Merlin 1D, you would need only 5 to 6 of the engines for liftoff. This makes the dry mass even lighter, which can go to further payload.
The engines combustion chamber and turbopumps will not have to be changed, only the nozzle. From images of the Merlin 1C it looks like the nozzle is bolted on. So you would bolt on a shorter nozzle, like you bolt on a longer addition when you producing the Merlin Vacuum.
2.)Weight of added reentry/landing systems.
a.)Thermal protection.
Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):
Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:
Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Command/Service_Module#Specifications
However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:
Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shuttle_thermal_protection_system#Weight_considerations
Also, for the X-37b the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:
X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/sis/x37b_otv/x37b_otv.html
For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:
Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index.php?topic=23522.msg754168#msg754168
while being able to survive lunar and even Martian reentry speeds.
SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.
Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as the high temperature-resistant aerospike to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
b.)Weight of the wings and the landing gear.
A common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:
The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downloads/JBIS_v57_22-32.pdf
On that same page the landing gear weight is the only 1.5% of the take-off weight.
For a vertical take-off vehicle these low weight proportions should apply to the dry weight.
3.)Increasing Isp with altitude compensation.
The most important characteristic for having a high vacuum Isp for an engine is having a long nozzle, i.e., area ratio, much more so than the chamber pressure. For instance the RL-10B2 while having a combustion chamber pressure of only 40 bar, has a vacuum Isp ca. 465 s. This comes from its very long vacuum optimized nozzle. And the Merlin Vacuum increases its vacuum Isp to 342 s from the 304 s of the Merlin 1C just by having a long nozzle extension attached.
The problem with this for a SSTO is that long nozzles give poor performance at sea level. In fact it's even worse then this. A vacuum optimized nozzle used at sea level can cause instabilities in the exhaust that can damage or destroy the engine. So the purpose of the altitude compensating nozzle is they can give this high performance of vacuum optimized nozzles, yet still work efficiently as well at sea level.
4.)Saving weight and increasing payload with a next-gen shuttle.
My comment about the shuttle orbiter being too heavy for high payload delivery was only in regards to the fact that every extra pound you take to LEO has to be subtracted from your payload capacity. The shuttle orbiter at ca. 80,000 kg subtracts majorly from the payload capacity of such a large launcher. With the lightweight composites now available we can cut significantly into that orbiter dry mass and therefore increase our payload.
Bob Clark
