Overview of the LMR N-Prize Approach

The approach we're taking is one based on my personal experience and research; it's not the only way to accomplish the N-Prize, but in my opinion it is the most straightforward, most expedient, and lowest cost.

Some of you have already gathered high-level insight into our plans; even if you hadn't followed the original N-Prize forum, there is enough information in the FAQ (and interview) to provide a reasonable outline. For those who haven't yet read the FAQ, we'll cover the basics of what we're doing (and not doing) along with some additional detail.

First, a little about what we're not doing. We're not doing a balloon launch (rockoon), air launch, jet assisted booster, space plane, light gas gun, electromagnetic mass driver, space elevator, nuclear rocket, beamed power-craft, anti-gravity system, teleporter, or any of the more exotic launch concepts. Unfortunately, each has significant drawbacks (not to mention the shortage of known technology for some of them). Perhaps they will all work adequately someday, but even the most prosaic of these presents a greater difficulty (and a generally higher expense) than conventional rocket systems do today. So, we're using conventional rocket systems.

What kind of rocket systems? Well, we're not using solids (such as NASA likes to do for boosters; I've never been a huge fan of them for space apps, but they're fine for tactical missiles and the sort); we're not using hybrids (more on that in a later blog). We're not using liquid hydrogen (again as NASA likes to do – though I wouldn't mind using LH2 for the upper stages, it's just not feasible for this project); we're also not using any other difficult/exotic fuels or oxidizers.

So what are we doing? Our approach centers on a sub micro-scale, three(3) stage, ground/sea-launched rocket; this will be a serially staged design employing pressure-fed liquid engines and low-cost, easy to get propellants.

In my opinion, LOX is the best, safest, and easiest to use oxidizer available – and it's easily obtainable; WFNA is the next best. If the filling logistics can be solved, we'll use LOX for all stages - if not, we'll use WFNA in the upper stages and LOX for the booster. There are also a few alternative storables that are interesting but unlikely to merit serious consideration. Fuels are most likely hydrocarbon (probably a light HC such as propane or butane, but also possibly kerosene, diesel or heavier – all deliver fairly similar overall performance with different flow and handling characteristics). Alcohol and ammonia are remote possibilities as well (we've experimented with all of them).

Each of our three(3) stages employs a single engine, with two(2)-axis control (gimbaled) – nothing fancy; no clustering. The system is fairly devoid of kludges: no variable thrust differential steering, exhaust vanes, or gas injection schemes; roll control is handled via simple gas jets. All engines are low pressure, including the booster. The third stage engine is roughly 30 psi chamber pressure (or potentially lower), the second stage roughly 60 psi; the booster will run at ~200 psi (or otherwise generally between 150 psi and 250 psi depending on certain experimental factors). The targeted specific impulse for the booster is estimated to be >220 s (sl), >230 s nominally; upper stages are both estimated at >270 s (vacuum).

The entire rocket will be about 75 kg GLOW (or more generally between 50 kg and 100 kg). The ultimate weight depends upon the final parameters of the third stage (the prototypes of which are now at ~560 g (wet mass) including payload, with an approximate 0.8 mass fraction; refer to the LMR-A/S3-HG for an example of one of our experimental third stages*).

Based on this and extrapolated mass fractions, the second stage mass (stage only) is roughly 9 kg GLOW (w/ 0.8 overall MF inc. payload); booster (stage only) ~65 kg GLOW (w/ 0.6 overall MF inc. payload) – these are close estimates, but approximates. Structural mass (no payload) of the booster is about 20 kg, second stage roughly 1.35 kg, and third stage roughly 90 g. Propellant breakdown: Stage 1: ~45 kg; Stage 2: ~7.65 kg; Stage 3: ~0.452 kg – total propellant ~53.102 kg (includes residuals).

The upper stage propellant tanks will be comprised of aluminum beverage cans (refer to the LMR-A/S3-HG overview for one prototypical embodiment*). I've been using such cans for propellant tanks in rocket projects for over two decades and have been continuously searching for an opportunity to use them in an upper stage or micro-spacecraft (a few years ago I applied for a patent related to this). They are lightweight, well-built, pressure bearing, easy to get, and inexpensive (there are also numerous other common containers that make excellent propellant tanks). These beverage can tanks (along with the exit cones) are the most fragile structures in the rocket, and there are no minimum gauge issues. As a note, future upper stages can comprise very lightweight electroformed structures, yielding even better mass fractions.

The booster fuselage will be made from aluminum irrigation pipe which will double as the propellant tank; such pipe is fairly lightweight yet strong enough for handling both aerodynamic loads and the required tank pressures; it again presents no minimum gauge issues. The upper part of the tube serves as a receptacle for the bottom half of the second stage as well as for first stage pressurant tanks and chutes, plus the nosecone attachment area. A small portion of the second stage serves as a receptacle for the third stage; all staging is fire-in-the-hole strategy.

Launching eastward (+ ~400 m/s delta V) from either a ground-based or sea-based platform (if sea-based off the Southeastern U.S. coast), the first stage (T/W: ~2:1, up to ~8:1) will carry the upper stages in near vertical ascent. The booster (depending on performance) will thrust nominally until roughly 10-15 km, at which point it will throttle back to a low value (just enough to maintain a small net acceleration); this allows the rocket system to maintain acceleration into near vacuum conditions at ~50 km (or possibly somewhat higher) and helps alleviate propellant settling maneuvers; any significant steering (other than roll orientation) will be (ideally) above 10 km, or more preferably towards the latter part of the first stage burn (if reasonable given performance metrics); steering will comprise a slow pitch maneuver and is handled by our proprietary ACS/GNC systems.

Protection of the thin-structured upper stages is via a two-piece clam-shell nosecone; this will cover both upper stages and open several seconds before first stage cutoff; the second stage will ignite slightly prior to first stage cutoff (@ ~50-60 km; T+ ~106 s). Depending on booster performance and final trajectory, total velocity at stage separation will be ~800-1200 m/s; horizontal velocity will be ~600-1000 m/s (including ~400 m/s Earth boost). The ~20 kg (dry mass) first stage falls into the Atlantic via chute – it's non-toxic; recovery is not necessary (but possible).

The second stage (T/W ~4:1) will burn and continue to steer towards an Earth-relative horizontal attitude; it will impart ~3.5-4 km/s delta_V and will be have reached horizontal attitude mid-to-late burn. The third stage will fire just prior to second stage burnout; this will occur @ ~80-130 km (depending on booster performance and final trajectory) @ T+ ~160 s (burn time ~54 seconds); after separation, the lightweight second stage falls and burns up in the atmosphere over the Atlantic.

The third stage (T/W ~0.75:1) will continue to burn, steering towards and maintaining horizontal attitude until target velocity (~7736 m/s) is reached at the target altitude (~280 km); the engine will shut down upon reaching the target velocity at roughly T+ ~420 s (burn time ~250 seconds); third stage propellant remainder at engine shutdown estimated at ~50 g (>10%); this allows for longer burns and/or potentially higher orbits. 

A few seconds after the third stage burn is complete, the control system will release (via electromechanical actuation) the ~18 g, 1.0” LMR satellite (LMRSAT-A/E1) (which will have started transmitting during the launch sequence). The third stage will decay into the dense upper atmosphere and subsequently burn up; total time from launch to satellite deployment is roughly 425 seconds.

The satellite will remain in orbit until the upper atmospheric drag overcomes its inertia, bringing it down to also burn up, which will be >2 days (battery life is ~3 days). During that time it will transmit identification information via an approximate 250 mW pulsed (~1-2 second interval) spread spectrum transmitter at roughly 2.45 GHz center frequency (anticipated, but subject to change based on several factors). Planned identification will include signature, temperature, voltage, velocity, and ox/fuel mass.

Electronics are lightweight and straightforward; the control system is a unique micro-controller based system – it integrates with my (proprietary) ACS/GNC control and simulation software suite. Operation of this will be the subject of another post, but it greatly simplifies the overall guidance and control system; the sensors and guidance components are confidential, but are simple, small and lightweight.

Although there may be some potential alternatives and course corrections, it is not anticipated that there will be major changes to our plans. Still, the numbers presented here are estimates, and though arrived at via detailed calculations and simulations, nothing should suggest anything final. The information is to provide some visibility into our current LMR launch system approach; everything is subject to change based on future experimentation and/or research.

~Sage

*LMR-A/S3-HG overview slide available with $20 donation.

Revision History

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2010-06-03:

Minor adj to third stage burn time and altitude (orig. ~440 s; ~220 km); inc. estimate of third stage propellant remaining at engine shutdown