Rocket Fuel & Oxidizer whats the best combinations

SpaceNut · 2014-10-26 18:01:34

This if once out of the R&D phase seems to be the solution to the Big Dumb Booster possibly.

Los Alamos National Laboratory scientists
Novel rocket design flight tested

Scientists recently flight tested a new rocket design that includes a high-energy fuel and a motor design that also delivers a high degree of safety.

After years of development and bench-top static tests, the new rocket design was recently flight tested at the Energetic Materials Research and Testing Center's Socorro launch site, part of New Mexico Tech. The new rocket design was tested against conventional, high-energy commercial rockets to enable a comparison of data gathered on velocity, altitude, burn rate, and other parameters.
Researchers will now work to scale-up the design, as well as explore miniaturization of the system, in order to exploit all potential applications that would require high-energy, high-velocity, and correspondingly high safety margins.

Last edited by SpaceNut (2014-10-26 18:09:07)

RobS · 2014-10-26 18:41:08

Too bad they didn't say what the new fuel is! I doubt it's any sort of breakthrough.

By the way, Elon Musk gave a 1 1/3 hour interview at MIT last week, I think Thursday. If you google "Elon Musk MIT" you can find the link in one of the news items. He describes the 300 by 170 foot floating platform he will land the Falcons on, starting in December. He said there's a 90% chance that he'll refly a first stage next year, and that they'll be making 12 launches in 2015. He also said that he won't try to make the Falcon 9 second stage reusable, because kerosene/LOX doesn't have a high enough Isp. That task will be reserved for the methane/oxygen rocket he'll design and build next. He said that that would be 5 or 6 years from now.

RobertDyck · 2014-10-26 22:49:15

I also noticed they didn't specify the fuel. So I did a google search. Wikipedia has this...

Double-base (DB) propellants
DB propellants are composed of two monopropellant fuel components where one typically acts as a high-energy (yet unstable) monopropellant and the other acts as a lower-energy stabilizing (and gelling) monopropellant. In typical circumstances, nitroglycerin is dissolved in a nitrocellulose gel and solidified with additives. DB propellants are implemented in applications where minimal smoke is required yet medium-high performance (Isp of roughly 235 s) is required. The addition of metal fuels (such as aluminum) can increase the performance (around 250 s), though metal oxide nucleation in the exhaust can turn the smoke opaque.
Composite propellants
A powdered oxidizer and powdered metal fuel are intimately mixed and immobilized with a rubbery binder (that also acts as a fuel). Composite propellants are often either ammonium nitrate-based (ANCP) or ammonium perchlorate-based (APCP). Ammonium nitrate composite propellant often uses magnesium and/or aluminum as fuel and delivers medium performance (Isp of about 210 s) whereas Ammonium Perchlorate Composite Propellant often uses aluminum fuel and delivers high performance (vacuum Isp up to 296 s with a single piece nozzle or 304 s with a high area ratio telescoping nozzle).[8] Composite propellants are cast, and retain their shape after the rubber binder, such as Hydroxyl-terminated polybutadiene (HTPB), cross-links (solidifies) with the aid of a curative additive. Because of its high performance, moderate ease of manufacturing, and moderate cost, APCP finds widespread use in space rockets, military rockets, hobby and amateur rockets, whereas cheaper and less efficient ANCP finds use in amateur rocketry and gas generators. Ammonium dinitramide, NH4N(NO2)2, is being considered as a 1-to-1 chlorine-free substitute for ammonium perchlorate in composite propellants. Unlike ammonium nitrate, ADN can be substituted for AP without a loss in motor performance.
In 2009, a group succeeded in creating a propellant of water and nanoaluminum (ALICE).
The Constellation Program uses a mix of aluminum, ammonium perchlorate, a polymer of polybutadiene and acrylonitrile, epoxy and iron oxide.
High-energy composite (HEC) propellants
Typical HEC propellants start with a standard composite propellant mixture (such as APCP) and add a high-energy explosive to the mix. This extra component usually is in the form of small crystals of RDX or HMX, both of which have higher energy than ammonium perchlorate. Despite a modest increase in specific impulse, implementation is limited due to the increased hazards of the high-explosive additives.
Composite modified double base propellants
Composite modified double base propellants start with a nitrocellulose/nitroglycerin double base propellant as a binder and add solids (typically ammonium perchlorate and powdered aluminum) normally used in composite propellants. The ammonium perchlorate makes up the oxygen deficit introduced by using nitrocellulose, improving the overall specific impulse. The aluminum also improves specific impulse as well as combustion stability. High performing propellants such as NEPE-75 used in Trident II D-5, replace most of the AP with HMX, further increasing specific impulse. The mixing of composite and double base propellant ingredients has become so common as to blur the functional definition of double base propellants.

The Los Alamos article uses the term "high-energy".

Tom Kalbfus · 2014-10-27 09:13:44

How does it get us into space?

GW Johnson · 2014-10-27 09:30:56

If I had to guess, I'd guess that the Los Alamos rocket was a hybrid. They used a new "energetic" fuel and a liquid oxidizer (likely LOX). That's how you achieve safety by separating fuel and oxidizer. Pressure-fed LOX eliminates turbopumps, too, at the expense of a heavy tank. The fuel probably resembled an AP composite solid propellant with the AP left out. Might or might not have had some HMX or RDX in it. Military composite solid propellants have used those as minor additives for decades now.

As for getting into space, there's nothing wrong with using solid or hybrid boosters off the pad into the first part of the trajectory. The issue early on isn't Isp, it's raw thrust, and solids are really good for that. You add them to a two-stage liquid core, and stage them off before the first stage liquid core stages off. It works really good. This is what Atlas-5 and Delta-4 really are. It's also what SLS is supposed to be. There's a reason for that: because it works pretty good.

There's a square-cube scaling law at work here because real materials are only so strong. The shuttle SRB's have given solids a bad name because of a bad joint design made worse by its putative "fix", and because they were so big as to get dinged up too badly upon ocean impact way too often. Smaller SRB's survive better, and if the joints are done right, are less susceptible to failure and less susceptible to ocean impact damage. ULA just throws them away, but they really could be reused.

Reusing the liquid core first stage is what Spacex is pioneering. That requires a 10,000 fps reentry survival. As long as the stage tumbles, airloads will crush it. Under control, they can be landed.

Liquid core second stages are an entirely different matter: a 26,000 fps reentry. Entry heating is an even bigger issue than airloads. That one may or may not be solved anytime soon.

GW

RobertDyck · 2014-10-27 13:08:27

Don't put any sort of an SRB joint beside a liquid hydrogen tank. If you want to use solids, use a single piece tube that has no joints. If that means smaller solids, so they fit within a railroad tunnel, then so be it.

GW Johnson · 2014-10-29 09:27:55

I know of no tunnels on the rail lines from Salt Lake City east toward Florida. The rocket whose dimensions were set by rail tunnel passage was the German WW2 V-2, not the shuttle SRB's. That's misinformation off the internet, which is notorious for that. The V-2's were shipped by rail from the underground Mittelwerk in the mountains of Bavaria. There's several tunnels on those rail lines.

The shuttle SRB segment size was set by propellant mix size, pure and simple.

There is (or at least was) another solid motor plant on the Mississippi River, operated by what was then UTC-CSD. Its products were barged to the Cape. Mix size set similar segment size limits for its products, too. The largest-diameter solids I ever heard of were 140 inches, not shuttle SRB's 120 inches. All such motors are segmented due to mix size.

All of these manufacturing and mix size limits would apply to hybrids as well. It's just a different world from the liquid rocket business.

As for segment joint design, solid (and hybrid) flames are far "dirtier" with soots and solid/molten oxides than anything the liquid boys ever even thought of, much less had any experience with. Liquid design experience simply does not apply to the solids arena. You use one and only one O-ring seal at such joints. If it passes a low-pressure (3-5 psi) whole-motor leak check test, it will hold at 1000's of psi. There's 70+ years' solid motor manufacturing experience supporting that idea.

Everything about the shuttle SRB joints was wrong from the very beginning, due to NASA's liquids-only experience being misapplied to the design decisions on a solid, and NASA's response to the Challenger disaster only made it worse. The only reason they never had another segment joint failure was they never flew soaked-out that cold again. I could go on and on about NASA's incompetence with solids, but I won't.

Solids are what we currently know how to build in very large sizes. A useful NASA program would be to scale up the hydrids to similar large sizes. That way, the abortable hybrid could replace the solid as an even safer booster (and solids can be quite safe in and of themselves). But agencies and corporations resist change. It's a profit-paranoia thing for the companies. It's a fearful-amount-of-incompetence thing for the agency.

GW

Last edited by GW Johnson (2014-10-29 09:32:02)

SpaceNut · 2015-01-21 20:30:39

While searching for data on the New F-1B possibility I got this infor mation to share.

http://arstechnica.com/science/2013/04/ … -thrust/2/

The specific impulse [of the liquid hydrogen and liquid oxygen-powered RS-25 Space Shuttle Main Engines] at sea-level (lift off) conditions is slightly over 365 seconds...the engine produces about 365 pounds of thrust at sea-level for each pound of hydrogen and oxygen burned together each second. The gee-whiz part is the engine burns propellant at an extremely large rate, just under 1100 pounds (over half a ton) each second, so each engine produces around 400,000 pounds thrust (force) at sea-level."
The numbers are different with RP-1. "The best demonstrated Isp performance for hydrogen is almost 365 seconds and kerosene is 311 seconds," he went on. "So, if we were to design our two engines to the same thrust level we would see that the hydrogen fueled engine is about 17% (1.17 times) better at producing thrust per unit of mass flow into the engine. That means if both engine cycles were sized for the same thrust, the more efficient hydrogen engine would use 17% less mass in propellants to push on the vehicle with the same force."
This brings us back to the question of density versus efficiency. A 17% bump in efficiency and decrease in mass is a big deal—individual kilograms count when dealing with rockets. But that more efficient fuel takes up a lot more space, and Coates outlined that trade-off very clearly. "Liquid hydrogen has a density of about 4.3 pounds per cubic foot, or to put it differently, each gallon of liquid hydrogen only weighs a little over a half pound. A gallon of water weighs in at about 8.3 pounds. Kerosene fuel is much more dense than hydrogen at about 50 pounds per cubic foot or just over 6.7 pounds per gallon. The kerosene fuel is well over 1100% (11 times) more dense than hydrogen fuel."
The question, though, is whether or not the practical side of the equation can balance the romantic. The Advanced Booster competition will run through 2015, at which point a winner will be chosen, solid or liquid. The F-1B could be the engine sending astronauts to Mars—or it could wind up as one more Wikipedia footnote.

kbd512 · 2015-01-22 16:39:54

If you do the math on the efficiency of LOX/LH2 and APCP, or pretty much anything else, it all looks remarkably similar to LOX/RP-1 in the end.

Chemical rocket technology is extremely limited and no amount of cleverness is going to significantly improve efficiency or thrust-to-weight beyond what's already been achieved.

The complication of high pressure turbo machinery and fabrication of parts from exotic alloys more or less guarantees that any rocket using that type of technology is going to be prohibitively expensive.

A large, pressure fed rocket is simply too inexpensive and not enough of an engineering challenge to interest NASA or its contractors. If a rocket only costs a few million to fabricate, doesn't require much in the way of technology to fabricate, and can be launched anywhere there's water, that's never something you'll see them do because there won't be hundreds of millions in profits from cost plus contracts.

The F-1 was a remarkable achievement for its day, but there are simpler and more cost effective options for heavy lift. Nobody is even trying because there's not enough money in it. FYI, F-1B boosters for SLS are never going to happen because the MLS would have to be redesigned, there's no room for the TSM's on the pad, and the attendant storage tanks, pipes, and pumps for the RP-1 would have to be built.

If the spectacular reduction in parts count and less expensive methods of fabricating the major assemblies make the F-1B inexpensive enough, there might still be a market for the engine, irrespective of whether or not SLS uses it.

GW Johnson · 2015-01-22 21:35:01

Here’s why I think the kind of Isp comparisons we typically see are often deceptive, because the conditions and assumptions are unstated. Doing the basic ballistics gets you the real comparisons. The measure most independent of nozzle conditions is chamber characteristic velocity c*, which does still depend weakly upon chamber pressure, so it needs to be stated when quoting a number.

Here are some 1000 psia chamber pressure values for c*, and the oxidizer/fuel ratio r that goes with it. These r’s are not stoichiometric, they are always a little fuel-rich. Both r and c* are weak functions of chamber pressure Pc. A good empirical correlation is c* = k Pc^ m, where typically m is a small-number exponent on the order of 0.01. It is entirely empirical, and determined by tests. It is always larger than the value you would estimate from thermochemical calculations made at different chamber pressures.

Oxidizer fuel c*, fps r
O2 H2 7950 4.0
O2 RP-1 5900 2.55
O2 NH3 5880 1.41
N2O4 UDMH 5680 2.6
HNO3 RP-1 5180 5.0

All else being equal (which is a lot to worry about), the 1000 psia c* (and thus Isp) of O2/H2 is about 134.7% that of O2/RP-1. That’s for the same Pc, the same ambient backpressure Pamb, and the same best-expansion nozzles at the two slightly-different specific heat ratios.

What you do with this kind of data is use your ratio Pc/Pamb and the specific heat ratio of the gases to determine your thrust coefficient. That has to include a nozzle kinetic energy efficiency (which is half-angle dependent) that applies to the mVe term in thrust, but not the (Pe – Pamb)Ae term. These calculations are fairly standard in the textbooks. The trickiest part is incorporating correctly the nozzle kinetic energy efficiency. It’s often left out of textbook equations.

Once you know your thrust coefficient at your design chamber pressure, and at your design ambient pressure (at which thrust is to be rated), then you can size the engine: F = Pc At CF. Then use the c* to set propellant flow rate: w = Pc At gc / c*. Then finally the Isp is F/w at that point.

For other operating points at different ambient backpressures, you have to refigure CF, then use the thrust equation F = Pc At CF to find that thrust, and the flow rate w = Pc At gc / c* is unchanged. Isp then changes. What you don’t want is an overexpanded exit nozzle where Pe < Pamb, because the (Pe – Pamb)Ae term in thrust goes negative. Overexpand enough, and the nozzle flow shock-separates, so that thrust is drastically reduced. There are empirical correlations for this, but precision isn’t possible.

Adding throttle-down to the design complicates this further, and reduces performance, simply because Pc is lower when you throttle down. That limits your expansion and thus your CF, to much lower values, which in turn lowers Isp. You may have to tolerate some overexpansion, but you need to stay far away from pressure ratios that would risk separation.

All these things affect thrust and Isp drastically. Chamber pressure affects c* weakly. So comparing c*’s is the more reliable means to evaluate propellant combinations. Thermochemical codes often report theoretical c* among all the other results. Rocket c* efficiencies are usually pretty high, especially in the larger sizes.

GW

Mars_B4_Moon · 2024-03-22 11:55:34

Why are nitric acid and hydrogen combinations not used as rocket fuel?
https://space.stackexchange.com/questio … ocket-fuel

New Mars Forums

Announcement

#1 2014-10-26 18:01:34

Rocket Fuel & Oxidizer whats the best combinations

#2 2014-10-26 18:41:08

Re: Rocket Fuel & Oxidizer whats the best combinations

#3 2014-10-26 22:49:15

Re: Rocket Fuel & Oxidizer whats the best combinations

#4 2014-10-27 09:13:44

Re: Rocket Fuel & Oxidizer whats the best combinations

#5 2014-10-27 09:30:56

Re: Rocket Fuel & Oxidizer whats the best combinations

#6 2014-10-27 13:08:27

Re: Rocket Fuel & Oxidizer whats the best combinations

#7 2014-10-29 09:27:55

Re: Rocket Fuel & Oxidizer whats the best combinations

#8 2015-01-21 20:30:39

Re: Rocket Fuel & Oxidizer whats the best combinations

#9 2015-01-22 16:39:54

Re: Rocket Fuel & Oxidizer whats the best combinations

#10 2015-01-22 21:35:01

Re: Rocket Fuel & Oxidizer whats the best combinations

#11 2024-03-22 11:55:34

Re: Rocket Fuel & Oxidizer whats the best combinations

Board footer