Space mining company developing nuclear reactor and more for moon projects
https://www.space.com/space-mining-comp … ew-feustel
Atomic rockets are back
https://www.aerosociety.com/news/atomic … -are-back/
NASA’s Building a Nuclear Rocket That Would Get Us to Mars in Just 6 Weeks
https://singularityhub.com/2023/08/07/n … t-6-weeks/
My ancient Pratt&Whitney handbook lists c* for NTO-hydrazine as 5860 ft/sec at 1000 psia versus 5740 ft/sec at 100 psia. Small systems are almost always pressure-fed. So if you figure a vacuum thrust coefficient for whatever expansion you can achieve, then Isp = CF c* / gc. Pick gc to make units consistent. If they already are, then gc = 1.
Same ancient handbook shows a sea level Isp = 292 sec at 1000 psia chamber, and a vacuum Isp = 342 sec at 100 psia. I don't think those are properly corrected for nozzle KE efficiency, though. The vac numbers are for expansion area ratio 40. The sea level numbers are for expanded pressure equal to sea level backpressure (perfect expansion).
Nozzle KE efficiency is typically very close to 0.983, but it only applies to the mdot V term, not the P A terms. If you have an off-angle nozzle orientation, you must use its cosine to knock down thrust and Isp. That applies to all the terms.
GW
]]>Kbd512 thanks for the links
]]>Obviously this is very high level, but the following document explains a bit about NTO/MMH tank fabrication and testing:
Design and Manufacture of a Propellant Tank Assembly
This explains the method ESA developed to decontaminate Rosetta's MMH tank assemblies without damaging them:
]]>What governs propellant flow rate from the tank in a pressure-fed system is pressure drop across the tank outlet fitting. The outlet side of that is set by engine chamber pressure. That means you maintain a high CONSTANT pressure in the propellant tank for a constant flow rate, constant thrust burn.
To do that means you have a high-pressure gas reservoir fed to the propellant tank through a pressure regulator. For a 2000 psig engine, the tank might be 3000 psig. The regulator must keep that 3000 psig on the tank as the gas reservoir pressure drops. It takes a positive pressure drop across that regulator to flow gas through it at all.
Just hazarding realistic guesses, say your reservoir operates between 8000 and 4000 psig. The reservoir volume must be such that gas mass expelled between those pressures equals the mass that fills the propellant tank as it empties at a constant 3000 psig. It's not a trivial gas bottle volume in relation to the propellant tank volume. You minimize gas bottle volume by going to very high initial gas pressures indeed, which does make that bottle heavier.
L* = chamber volume/nozzle throat area = 100+ inches for most propellants. Determined initially by von Braun at Penemunde for LOX-ethanol in his A-4/V-2 work. Varies by propellant combination, but not all that much.
GW
]]>It's easy enough to have sufficient pressure when the tank is full. But you need to pump in enough gas to keep pressure up until the tank is empty. Say you'll use the tank until 5% of the fuel is left. In that case, your pressurant gas has expanded by a factor of 20, and has dropped pressure by a factor of 20. If the pressure at that point is 1 bar, you need to start at 20 bar. That means you need a very heavy tank: 20 times heavier than a tank that's good for 1 bar. Its also why its heated to make it expand once out of the tank that its stored in.
http://space.stackexchange.com/question … tabilities
To be theoretical the fuel residence time in a combustion chamber is given by the Characteristic length(usually denoted by L*)(minimum length that the fuel will remain in the combustion chamber and nozzle for complete combustion to take place)
L∗=q∗V∗ts/A
q is the propellant mass flow rate, V is the average specific volume, and ts
is the propellant stay-time A is the sonic throat area
High c* efficiency correlates with long chamber length (L/D > 1+) and larger contraction area ratios (Ach/At > 4), both of which raise the L* parameter. Once you get to an acceptable L*, c*-efficiency should be above 90%. In a good design with well-designed injection spray patterns and with long chamber and large contraction ratio, c* efficiency is above 98%.
Isp needs to figure on delivered thrust divided by the actual propellant flow rate. That propellant flow rate is the ideal value in the final chamber feeding the nozzle, plus whatever is bled off upstream to feed the turbopumps.
For a pressure-fed design, there is only one chamber and it has a chamber pressure. All the flow rate is ideal, so wdot = Pc At g / c* is all the massflow that there is from the tankage, while thrust F = CF Pc At. In that case (and only that case) Isp = CF c* / g.
Even for pressure-fed designs, you still have to worry about selecting chamber length and nozzle contraction, and getting good spray patterns, to achieve high c* efficiencies. I would use L/D around 1.5 and Ach/At exceeding 4 to get efficiencies at 99+%, given good spray patterns. Most violate this, and so most have c* efficiency losses. It's a tradeoff of achievable performance versus the weight of the components you design.
In pressure-fed designs, the old 1950's - 1960's rule of thumb was you need a feed pressure / chamber pressure ratio near 2 to achieve combustion stability. They tell me that ratio is closer to 1.2 today. There's a similar ratio across whatever kinds of valves control the propellant feed rate for a given pressure ratio. My best guess is a tank pressure around 1.4-1.5 times whatever chamber pressure you end up designing to. That's just a rough figure.
There's just no way to make pressure-fed propellant tankage really lightweight. For a 2000 psig rocket chamber, you need something like a 3000 psig tank. The modified Barlow's formula (using vessel ID) relates hoop stress to vessel pressure, which in turn sets a lower bound on your tank weight: stress x 2 x twall = Pressure x ID.
You simply cannot use material ultimate stress capability in a practical design. Even a 1-shot design should stay under material yield stress. The hotter you let the materials get, the lower those allowable-stress numbers get. Reused designs need to set an allowable stress about 10 times less than the S-N curve levels indicate (and that's before you knock down for temperature), for the appropriate load case in Mil Handbook 5, and an N defined by the expected life. If the structure carries other loads as well (most do), then those may (or may not) actually set your design. Every situation is different.
It does get quite complicated. Sorry, that's just life.
Titanium won't save you. It's hard to fabricate with, and it's no better hot than plain carbon steel (pretty much crap above 750 F). Its only advantage is that it's just about half the weight of steel. Austenitic and martensitic stainless steels will go a lot hotter, most near 1200 F. Don't even think about aluminum. It's crap at 300 F.
There are a couple of austenitic stainless alloys usable even hotter (near 1700 F), but they are quite a bit weaker there than at 1200 F, where they are just about the same strength as the others. The only real difference is freedom-from-scaling to higher temperatures, not strength. 316 and 310 are the ones. Softened-butter weak, but free of scale.
Austenitics (300-series) must be used at low annealed strength. They do work-harden, biut modest heating anneals them again. Martensitics (D6ac and the 4000-series, and T-250 and similar) can be heat-treated to high strength, but there are severe exposure limits above which they revert to low annealed strength (usually under 1200 F). Scaling and corrosion get very bad beyond 1200 F with nearly all of these.
There are some super alloys (that really aren't steels), that will go hotter still, but they are super-expensive, and very difficult to fabricate. You find them in combustor cans, turbine blades, and afterburner components.
GW
]]>Can you explain how pressurization in pressure fed systems relates to engine cycle losses or other issues affecting accurate determination of appropriate expansion ratios and Isp?
]]>The same thermochem code tells you what your gas specific heat ratio is (gamma). You need that to determine thrust coefficient at your chamber pressure and whatever backpressure exists. It is a function of that backpressure, coming from F = mdot V + (Pexpanded - Pback)Aexit. Thrust maximizes when Pexit = Pback, obviously. As you climb in altitude, Pback reduces. Pexpanded is a fixed ratio to whatever chamber pressure you are using.
When Pexpanded - Pback is positive, thrust is higher for the same wdot V term, but you could have expanded further (to higher V) to actually optimize things. We call that condition underexpanded. When Pexpanded - Pback is negative, thrust is reduced for the same wdot V term. You need less expansion ratio to increase thrust back to optimal, but it will cost you V. We call that condition overexpanded.
But there is a max Pback limit that causes shockdown flow separation in the exit bell (which just absolutely kills your thrust!). That can be approximated, but not predicted precisely. It is a risk if you are too overexpanded. Being underexpanded is no risk, it just means you get a little less thrust than you could possibly have. Which is why launch vehicle engines have nozzles sized for sea level perfect expansion and max thrust (when the weight to be lifted is maximum), and then operate increasingly underexpanded as the vehicle rises. Thrust increases as they climb, but not all the way to ideal values.
Nozzles that operate in vacuum have much larger expansion bells for much higher V. You cannot expand all the way down to vacuum as the exit Pexpanded, because the device would be of infinite size! You must fit within physical dimension limits of your vehicle. There are NO "perfect-vacuum-expansion" designs, only designs that fit certain dimensional constraints.
Tables and charts exist which have thrust coefficients CF reflecting these phenomena. These usually include the nozzle kinetic energy efficiency (which reflects bell expansion half angle). They are usually available for this or that gamma (specific heat ratio). F = Pc At CF, where Pc = chamber pressure and At = throat area.
Without belaboring derivations, ideally CF c* /g = Isp, except for engine cycle effects! To get those, you calculate ideal chamber-nozzle wdot = Pc At g / c* (where g is the gravity constant that makes units consistent, not an actual acceleration of gravity). Isp = F/wdot as a definition, but you need to add some extra wdot to your ideal amount, to reflect the bled-off massflow that operates your propellant turbopumps. That's how your engine cycle gets into it, really. Turbopump bleeds inherently reduce Isp, because you must inherently flow more propellants than what creates the nozzle thrust forces.
Reported sea level Isp data usually does not reflect engine cycle losses, and usually does not have the nozzle kinetic energy efficiency figured into CF. Reported vacuum Isp data lacks cycle losses, lacks nozzle efficiency, and usually they never tell you what expansion ratio they really used, which is utterly crucial to CF. So beware of other people's Isp. It's usually one or another sort of lie.
Does that help? I hope?
GW
]]>Draco and SuperDraco both use NTO/MMH. This bi-propellant combination produces higher Isp than HAN monopropellant alone. The poor vacuum performance is probably due to under-expanded nozzles. The expansion ratio looks like it's optimized for throttling at sea level, probably because it was intended as a launch escape system. However, I'd get a real rocket scientist like GW to confirm that.
SuperDraco is regeneratively-cooled, so if 2 SuperDracos weigh less than 4 uprated AJ10-190, then my ascent stage could use SuperDracos instead of AJ10-190's. It would produce even more thrust (32,000lbf for 2 SuperDracos to 30,000lbf for 4 uprated AJ10-190's). With two properly expanded SuperDraco's producing a 316s or better Isp, assuming two SuperDracos weigh less than 400kg, the inert mass fraction of the tankage and feed system can rise to something inline with commonly used materials.
]]>SpaceNut,
The article says they're using MR-80B (3100N) and MR-107U (275N) thrusters. MR-80C (3100N) and MR-104D (440N) are thrusters I made up. They're notional variants of the MR-80B (3100N) and MR-104C (440N), both of which are real flight hardware, that have been modified to use materials compatible with the slightly acidic HAN (AF-M315E) monopropellant.
* HAN monopropellant thrusters do not use double valves as do thrusters using N2H4 / hydrazine (MR-80B and MR-104C are examples thereof) use. HAN requires high temperature catalyst beds (200C to 300C) compared to hydrazine (100C to 150C), and thus more electrical power, to decompose the propellant. Hydrazine will start to decompose purely as a function of contacting the catalyst bed, although Isp will be lower without pre-heating. The switch to single valves for HAN-fueled thrusters is principally a characteristic of the high heat requirement to facilitate propellant decomposition. Simply put, propellant leaking through the feed valve won't fire the thruster without a hot catalyst bed. HAN is also less likely to decompose in hot propellant feed lines connected directly to the thruster, compared to hydrazine. This simplifies the propellant feed system.
* HAN does not require continuous heating of the propellant tanks and feed lines. It has a glass transition temperature, but doesn't freeze like hydrazine will if it gets too cold during the deep space transit. Basically, you heat the tanks and feed lines when you intend to use the thrusters.
* AF-M315E (a trade name for a specific HydroxlAmmonium Nitrate or HAN blend / formulation developed by Air Force Research Laboratory or AFRL and AeroJet-Rocketdyne) monopropellant is comparatively non-toxic to humans, in relation to hydrazine. That is another reason for the simplified propellant feed system. Spills or leaks are not events that pose immediate danger to human health during propellant handling. For comparison, a gasoline spill is more hazardous to your health and a hydrazine spill would seriously injure or kill you without protection.
Further notes:
16. The Cygnus PCM has 4 variable length (extendible) landing legs attached to steel wire wheels that can caster. Each wheel hub has a small electric motor so the PCM can move at low speeds. This is how the Cygnus PCM is mated to the Earth return stage. The PCM drives over to the Earth return stage after landing, extends the legs to raise the PCM above the upper stage, and then rotates about its axis to lock the PCM in place.
17. The empty RCS module stays attached to Cygnus during ascent because it is later required during Earth return to stabilize the Cygnus PCM during Earth reentry. After the Cygnus PCM ascends to Mars orbit using the ascent stage, they have a series of three spacewalks to perform to prep their Cygnus PCM for Earth return.
Space Walk I - Transfer food and water from the Earth Return Stage to the Cygnus PCM
Space Walk II - Refill the empty HAN and GN2 tanks using tanks integrated into the Earth Return Stage
Space Walk III - Attach a new HIAD to the Cygnus PCM for Earth reentry
18. The Earth reentry HIAD is specifically designed for that purpose. All HIAD's have to be constructed specifically for their intended payload mass. It's the same basic technology set and materials, but different donut sizes, material thickness, or material layers. An Earth reentry HIAD may have five layers of thermally protective fabric whereas a Mars reentry HIAD may only have three layers.
19. If SLS and SEP can deliver Lockheed-Martin's Mars orbital station, better known as Mars Base Camp, to LMO, then the spacewalks are performed while the crew stay aboard the space station in Mars orbit. There will be 2 to 4 crews (4 to 8 astronauts) loading supplies at the same time, so an ISS node module with 4 docking ports and an ISS MPLM module to store consumables is the bare minimum required to service 4 Cygnus PCM's. An ISS node module and ISS Destiny (lab habitat) module is more realistic.
20. 1g artificial gravity is provided on the departure flight by tethering off to the expended upper stage. Artificial gravity may or may not be practical for Earth return. The PCM really needs to stay solidly attached to the Earth return stage because that stage is responsible for slowing the PCM just prior to reentry. If that fails, you get BBQ'd during reentry and that's not a good way to end an otherwise successful mission.
21. If NASA is willing to give up 30 to 60 days of surface time by spiraling into LMO using SEP and/or spiraling into LEO using SEP, then reentry velocities drop substantially. If that Mars orbital station ever materializes, it's a good place to inspect PCM's prior to committing to a reentry. If I was an astronaut, I'd want to determine whether or not my PCM, RCS, and HIAD modules were in good working order after my trip through interplanetary space.
22. Using SEP to spiral into both LMO and LEO, I eliminate the requirement to attach a second HIAD to the Cygnus PCM. The RCS tank still requires a refill. The food and water still require replenishment prior to Earth return. Upon Earth return, the PCM's dock at ISS for samples offload, crew de-briefing, and return to their respective countries along with astronauts from ISS. The Americans and Canadians return to America aboard Dragon V2, the Asians return to China aboard Shenzhou (assuming they participate; they seem to be more interested in the moon), and the Europeans and Russians return to Kazakhstan aboard PTK NP. There is no need to store NTO/MMH in the PCM service module or the Earth Return Stage, either. HAN and SEP are sufficient for all mission requirements except the Mars ascent. NTO/MMH, as dangerous as they are, are required by the Mars Ascent Stage to achieve sufficient Isp so that a single Falcon Heavy can deliver the stage to Mars.
GW,
IRVE-3 used rotation for stability and cold-gas RCS for attitude control at the entry interface. Spinning humans may not work so well. MSL used spin for stability and RCS for attitude control. Could spin negate or lessen the RCS requirement, or is that a bad idea?
A new method for HIAD control NASA is working on is called the Cable-Controlled Aeroshell Deceleration System (CCADS). It is intended to provide coarse and fine HIAD flight control during descent by varying L/D to adjust cross-range, a concept much like using the body flap on the orbiter to adjust aerodynamic lift (CCADS basically uses cables to "pinch" the HIAD):
CCADS Technical Paper Final Draft
The "HEART Flight Test Overview" report said shallow angles of attack (10 degrees or less) negated the requirement for thermal protection of the exposed portion of the Cygnus. Is that an error with the Thermal Desktop heating model used to determine peak heating or is that possible? The CCADS flight profile uses entry angles greater than 10 degrees. Is it possible to use a 10 degree or less followed by actuation of CCADS to increase lift and establish a more horizontal glide path to the ground?
The computer models indicate that HIAD (no parachutes or CCADS) is subsonic at 15km, but then velocity is nearly constant between .3km/s and .2km/s all the way to the ground. There is virtually no deceleration from .2km/s at approximately 5km. Apparently, high subsonic flight speeds are required to generate lift. CCADS lengthens flight duration considerably, permitting astronauts to actually "fly" Cygnus to a landing target, but it's either traveling at .2km/s or it ceases to generate lift and subsequently falls out of the sky.
A subsonic parachute and mortar system would weigh more than rocket engines and propellant with .25km/s dV. I give up. There's no better way to do this.
My proposed propulsion module is mounted atop Cygnus over the docking ring since HIAD is on the bottom. It has 8 MR-80C's (AeroJet-Rocketdyne MR-80B's modified to use HAN / AF-M315E monopropellant) for retro-propulsion and 16 MR-104D'S (AeroJet-Rocketdyne MR-104C's modified to use HAN / AF-M315E monopropellant) for attitude control. Isp (vac) for MR-80C and MR-104D is presumed to be 250s. AF-M315E's density is 1.47g/cm^3. A US gallon of AF-M315E weighs approximately 12.26774lbs, so 450kg of the stuff is slightly less than 81 gallons. My retro-propulsion engines are radially mounted over the docking ring in a sort of "nosecone" for the Cygnus PCM (a misnomer since Cygnus is traveling base forward throughout EDL).
Notes:
1. Cygnus won't fly aboard Antares, thus no requirement for the structure to withstand 8.5g peak acceleration, so it'll be a bit lighter
2. HIAD is separated just prior to retro-propulsion and Cygnus becomes 1,000kg lighter
3. RCS module is permanently attached to Cygnus, but hinged so the crew can ingress / egress through the PCM's top hatch
4. Cygnus is modified with the addition of a bottom hatch for normal ingress / egress on the surface of Mars
5. Solar panels for surface power are stowed inside Cygnus and manually attached by the crew to the top hatch
6. Cygnus landing gear consists of composite struts with small wheels containing electric hub motors for 2.5km/h max speed
7. Cygnus lands several kilometers from the ascent stage to avoid potentially damaging the ascent stage during retro-propulsion
8. Consumables mass assumes the crew consists of 1 average man and 1 average woman
9. Service Module provides power and propulsion for transit
10. Apart from providing transit power and mid-course correction burns, the primary purpose of the service module is to decelerate the PCM approximately 1km/s just prior to reentry, at which point in time it detaches from the PCM
11. My previous mass estimates for a scaled-up HIAD were used for the ascent stage. Mass was greater because the ascent stage has a greater mass. I mistakenly added that mass to the Cygnus reentry mass. My Cygnus PCM is within the 5600kg limit imposed by the HEART HIAD precursor mission design.
12. Cygnus is positioned atop the ascent stage
13. Ascent Stage uses 4 AeroJet-Rocketdyne AJ10-190's (the uprated OMS-E / OME engine developed for Orion); essentially a reprise of MIT's Scott Alan Geels Mars Ascent Vehicle, 4.2km/s dV (accounts for drag and gravity losses); requires a 10% or less inert mass fraction; nearly maxes out Falcon Heavy's throw capability, requiring 13t to TMI
14. Earth Return Stage uses 1 AeroJet-Rocketdyne AJ10-190 (the uprated OMS-E / OME engine developed for Orion)
15. Consumables and HIAD #2 for Earth return are delivered with the Earth Return Stage (some assembly required)Mass Estimates:
Service Module (1 AeroJet RocketDyne AJ10-190): 4,750kg
Gross Reentry Mass: 5,250kg
HIAD: 1,000kg
Cygnus PCM gross mass (reentry; includes RCS module): 4,250kg
Cygnus PCM (structure, avionics, life support - CAMRAS / MOXIE / IWP, batteries, solar arrays for surface power): 1,650kg
Consumables (food and water): 1,850kg
Astronauts (we're sending small people to Mars): 150kg
RCS Module: (450kg HAN, 68kg for 8 MR-80C, 30kg for 16 MR-104D, 52kg for structures): 600kgWell, there it is. It's quite minimalist, but it should get the job done. I would feel better about this if there was another Cygnus loaded with consumables and parked in the landing area. A Mars orbital station like Lockheed-Martin's Mars Base Camp would be nice, too. All the Falcon Heavy payloads range from 11t to 13t. If I had 15t to work with, the mission hardware elements could be more robust. Unfortunately, 13.6t is all Falcon Heavy can deliver. C'est la vie.
Affordable Human Exploration of Mars Program
MISSION SIZE: 4 Astronauts per mission
PRE-MISSION CONTENT:
Three Phased 5 Year Development and Demonstration Programs:
Phase I:
Deep Space Habitat (DSH) with Artificial Gravity (AG)
Closed-Loop Environmental Control and Life Support System (CL-ECLSS)
High Efficiency Solar Cell and Battery Systems
Passive Radiation Shielding System (PRSS)
Phase II:
SEP Cargo Transfer Vehicle (SEP-CTV)
SEP Integrated Transfer Vehicle (SEP-ITV)
Mars Surface Habitat (MSH)
Adaptable Deployable Entry and Placement Technology (ADEPT)
Phase III:
Trans-Mars Injection (TMI) and Trans-Earth Injection (TEI) Kick Stages
Mars Descent Vehicle (MDV) / Mars Ascent Vehicle (MAV)
Active Radiation Shielding System (ARSS)
Mechanical Counter-Pressure Space (MCP) Suits
LAUNCH:
Phase I Demonstrators:
DSH-1 (LEO)
DSH-2 (L1)
Phase 2 Demonstrators:
CTV-1 (150kW-class SEP 30t cargo transfer of 1 MSH using ADEPT for aerocapture and reentry)
Phase 3 Demonstrators:
CTV-2 (150kW-class SEP 40t cargo transfer of 1 TEI stage to LMO)
CTV-3 (150kW-class SEP 30t cargo transfer of 1 MAV to Mars using ADEPT for aerocapture and reentry)
DSH-3 (tele-robotically operated DSH-ITV equipped with ARSS transfers to Mars using TMI stage, captures in LMO, tele-robotically operated MAV simulates transfer of crew after completion of surface missions, CTV-2 mates TEI stage to DSH-3, transfers to Earth using TEI stage, and captures at L1)
TRANSIT TO MARS:
7 Month transfer to Mars using chemical kick stage followed by DSH-ITV spiral in to LMO
TRANSIT FROM MARS:
7 Month transfer to Earth using chemical kick stage followed by DSH-ITVE spiral in to L1
ENTRY, DESCENT, LANDING:
3t MDV (pressurized MAV without fuel tanks and engines for ascent) transferred to Mars using SEP-CTV, docked to DSH-ITV for crew transfer after DSH-ITV arrives in LMO, and landed using ADEPT / super-sonic parachute / retro-rocket
ASCENT:
15t pressurized MAV transferred to Mars using SEP-CTV and landed using ADEPT / super-sonic parachute / retro-rocket
RETURN:
MAV docked to DSH-ITV for crew transfer, TEI mated to DSH-ITV, capture at L1 using SEP
ENERGY AND LIFE SUPPORT:
DSH-ITV uses solar panels for power
DSH-ITV is equipped with next-generation CL-ECLSS, intended to replace current ECLSS on ISS, for life support
MSH uses solar panels for power
MSH equipped with next-generation CL-ECLSS, intended to replace current ECLSS on ISS, for life support
MSH equipped with oxygen generation demonstrator
MSH equipped with Martian regolith water extraction demonstrator
MISSION CONTENT:
Focuses on technology development that maximizes use of critical path technologies already identified by NASA for human exploration of Mars
Directs funding towards payload, rather than launch vehicle development
Sacrifices some measure of simplicity to avoid further launch vehicle development
Permits sustainability initiatives (ISRU and ISPP) for permanent habitability of Mars, but does not rely on sustainability to land humans on Mars
Intended to achieve the goal of landing humans on Mars for surface exploration, not colonization
Each major mission hardware component has a replacement as backup
Uses commercial launch services and capsules for affordability
COST/INCOME:
This plan requires present levels of funding for NASA. Each 5 year technology development phase spreads the cost of research, development, and demonstration as evenly as possible, given the logical horse/cart requirements for achieving the ultimate goal of landing humans on Mars.
I put the development phase costs in the $50B to $75B range, over the course of 15 years. At the high end, that's 25% of NASA's budget. I think it's reasonable for NASA to devote a quarter of their budget towards a real space exploration program. NASA has repeatedly said that Mars is the prize. I say it's long past time for them to put their money where their mouth is.
OTHER ASPECTS:
I think it's reasonable to orbit a constellation of advanced TDRS and GPS satellites over a planet we intend to make our own. Obviously there's quite a ways to go to establishing an infrastructure required to colonize Mars, but the first steps have to be taken and must be taken by our government.
Subsequent missions can land advanced pressurized rovers for mobile surface exploration and various sustainability experiments, but the first order of business is to simply prove that we can get there and come back. This mission architecture is intended to do that and nothing further.
louis,
If you require launch schedules that include launch vehicle and payload descriptions, mission objectives, etc, please let me know.
]]>