News: Why Not Space?

This author looks at some of the negative aspects of space travel and colonization. I think it is worth being a realist and considering both the benefits and problems in any proposition. This author has some cogent arguments. He could be right, he could be wrong but we should consider his ideas, I think.

Ed L

Why Not Space?
Do The Math, 12 October 2011
href="http://physics.ucsd.edu/do-the-math/2011/10/why-not-space

"In other words, I’m an insider—and a supporter. I whole-heartedly believe that space offers tremendous scientific promise." ...

"But I want to caution against harboring illusions of space as the answer to our collision course of growth on a finite planet. We live at a special time. We have enjoyed spending our inheritance of fossil fuels, and are feeling rather heady about our technological prowess. For many generations now, we have ridden an exponential growth track, conditioning ourselves to believe that our upward trajectory is an eternal constant of our existence."

Looking at a Solid First Stage Booster

Just because I thought it might be worth considering what an all-solids launcher

might look like using high-performance commercial ATK motors, I found what was a

reasonable booster suitable for the first stage. The following diagram shows the

numeric specifications and compares the solid first stage with an equivalent liquid stage. I’ve darkened out the parameters of the upper stages because those things haven’t changed from my last post.

If you look at the image below, one thing that stands out is that the solid booster is physically shorter (for the same diameter) compared to the liquid booster.

This is due to the solid booster's higher average propellant density, which is almost double that of the liquids that I had selected. In actuality, the liquid booster is lighter by about 200 lbs, but it is physically larger because of the lower propellant density.

A summarization of this Comparison

Now, the main impetus of this design experiment was to look at reasonable performance

solids as examples of upper stage vehicles and, then finally, as a first stage booster.

For amateurs to develop these kinds of vehicles, they would have to try to duplicate

the performance factors of these commercial motors with their own designed and fabricated motors. One good thing from this is that they have a known working example and explicit specifications for trying to develop their own versions of the commercial motors. This will simplify the development task somewhat.

Because I went with off-the-shelf motors, their specifications dictated many of the

requirements for the vehicle (such as stage delta V, thrust and mass ratios). In many ways, I consider these designs to be an upper limit of reasonable size for launching a payload as small as a 1/4 pound (113 grams). Think about it: a payload of 1/4 pound required a launcher with a GLOW of 1914 lbs. This is an effective payload:glow ratio of 0.00013 or 0.013 %. This is pretty poor overall. Nonetheless, with smaller launchers we'll see these kinds of ratio values. But we can do better. To do better means that a fundamental ratio between payload and total rocket takeoff mass will be a higher number. Let's examine how to get to smaller, more manageable rockets in upcoming future posts.

An All-Solid-Upper-Stages Booster

An All-Solid-Upper-Stages Booster

By Ed LeBouthillier

I don't want anyone to think that I'm picking on solids as being incapable for booster stages. To the contrary, they're very capable and offer many benefits of their own over liquid stages. To show their capability (and some aspects of their requirements), I'm going to do a quick analysis of an all-solid-upper-stages booster utilizing off-the-shelf solid rockets.

ATK [http://www.atk.com/corporateoverview/corpover_missiongroup.asp] produces some very high quality solid motors which are used in various military and space applications. Their 2008 catalog can be found online [

http://www.atk.com/capabilities_space/documents/atk_catalog_may_2008.pdf] so I thought I'd look at their offerings to look at what a solid-upper-stages booster might look like. Now, before you start saying "cool! let's do it," I should point out that these motors are what many of us would call "very expensive." But, they provide some ideas on what a small orbital launcher using solid upper stages might look like.

I've selected 3 ATK solid motors for the upper stages: the Star 5C, the Star 9 and the Star 15g as being representative of good solid motors suitable for this task.

This analysis has a number of caveats worth considering:

1. DELTA-V - I've only allowed for about 24500 fps (7467.6 m/s) for upper stages

2. CONTROL - I haven't included any weight for control systems (i.e. TVC)

3. GUIDANCE - No weight has been included for a guidance system

The idea in this analysis is merely to show what a small orbital vehicle might look like in a rough way.

Here are the results that I get:

	Stage 4	Stage 3	Stage 2	Stage 1
Oxidizer	AP	AP	AP	Lox
Fuel	HTPB/Al	HTPB/Al	HTPB/Al	Propane
	Star 5C	Star 9	Star 15G
Payload	0.250	10.110	51.111	257.707	lbs
OF Ratio	6.400	6.400	6.400	2.200
Oxidizer Density	121.700	121.700	121.700	71.23	lbs/cuft
Fuel Density	68.498	68.498	68.498	33.36	lbs/cuft
Avg Density	114.511	114.511	114.511	59.396	lbs/cuft
Average Isp	268.1	289.1	281.8	250	Seconds
Desired DeltaV	4928.1	9102	10674	10409	FPS
Body:Fuel Mass	1.2409	0.2853	0.1587	0.2467
Thrust	450	1200	1000	6350	lb-f
Payload Ratio	0.025	0.247	0.247	0.105
Structural Coef	0.554	0.222	0.137	0.198
Propellant Ratio	0.446	0.778	0.863	0.802
Mf/Me Ratio	1.771	2.661	3.245	3.648
Propellant Mass	4.400	31.900	178.300	1967.312	lbs
Oxidizer Mass	3.805	27.589	154.205	1352.527	lbs
Fuel Mass	0.595	4.311	24.095	614.785	lbs
Oxidizer Volume	0.031	0.227	1.267	18.988	cuft
Fuel Volume	0.009	0.063	0.352	18.429	cuft
Stage Weight	9.860	41.001	206.596	2452.648	lbs
MT	5.460	9.101	28.296	485.336	lbs
Me	5.710	19.211	79.407	743.042	lbs
Mf	10.110	51.111	257.707	2710.354	lbs
Max G's	78.811	62.464	12.593	8.546	g's
Cum Delta V	4928.050	14029.800	24703.410	35112.410

So, the Gross Lift Off Weight (GLOW) is about 2710 lbs (1011 kg). The payload experiences upwards of 79 g’s as the final stage is nearing empty (presuming a fairly constant near-average thrust which is what the thrust curve does show).

But, as a feasibility or model of a possible launcher, it shows what is possible. The first stage is a scaled-up Aerobee 150 sustainer. The body diameter is 20 inches, the finspan is 3 feet 11 inches and the total height is about 24 feet 2 inches.

A Short Survey of Off-The-Shelf Solid Motors for Orbital Upper Stages

A Short Survey of Off-The-Shelf Solid Motors for Orbital Upper Stages

By Ed LeBouthillier

Many suggest using off-the-shelf solid motors (or similar custom motors) for upper stages of small orbital launchers. In this discussion, I review some of the requirements for upper stages and the possibility of using solids as upper stage motors. I will presume that a basic 1/4 pound (113 gram) payload is selected.

OFF THE SHELF MOTORS

High Power rocketry uses motors that provide many benefits for someone considering upper stages for orbital launchers. They are efficient, use modern propellants, come in a wide range of impulses and require little development.

The following table lists a few different solid motors that might be suitable for upper stages and their parameters:

Manuf.	Model	Total Impulse (N-s)	Propellant Weight (grams)	Loaded Weight (grams)	Empty Weight (grams)	BP Ratio (λ)	SL Isp	Vac Isp
Aerotech	I305	450	302.1	581	278.9	0.92	150 s	211 s
Cesaroni	I303	538	270.0	500	230.0	0.85	189 s	235 s
Aerotech	I350R	2500	1400.0	2294	894.0	0.64	189 s	235 s

( Note: BP Ratio = Empty Weight/Propellant Weight )

Let me explain how I estimated the specific impulses for sea level (SL) and vacuum (Vac) exhaust pressures.

First, the sea level Isp is derived from the published data (and verified as being reasonable). The equation is:

                                    Avg Thrust * Burn Duration

            Ideal Isp =        --------------------------

                                         Propellant Mass

These propellants are specified as being composite propellants. I used Propep as the

combustion code to estimate the Isp. I put Ammonium Perchlorate and HTPB into ProPep. I then set the mixture ratio similar to what is published as commonly used. I presumed that the published value represents 90% of the theoretical maximum Isp. Therefore, if 150 seconds is the value derived from published figures, then the theoretical ideal value of the Isp is 150 seconds / 0.90 = 166 seconds. I then adjusted the chamber pressure in Propep until I got a theoretical value equal to this Ideal Isp. I then calculated the Isp in a vacuum using Propep and then multiplied that value by 90% to get the vacuum Isp. It’s rough, but it gives meaningful statistics for comparison.

IMPLICATIONS

Based on the above figures, we can estimate the likely delta V from one of these motors.

If we presume no payload, and just the motor weight, then using the Aerotech I350R as an example we have:

            dv = g * Isp * ln( Mf / Me )

            dv = 9.8 * 235 * ln( 2294 g / 894 g )

            dv = 2303 * ln( 2.57 )

            dv = 2303 * 0.94

            dv = 2164.82 m/s (7102 fps)

Since we need to provide about 7467.6 m/s (24500 fps) to 7772.4 m/s (25500 fps) in the upper stages, we would need about 7772 / 2165 = 4 stages at this performance level (for a total of 5 stages with a first stage). Presuming a 113 gram payload, a 113 gram guidance and control system for the 5th stage, we have:

	Stage 5	Stage 4	Stage 3	Stage 2	Stage 1
Oxidizer	AP	AP	AP	AP	Lox
Fuel	HTPB	HTPB	HTPB	HTPB	Propane
Payload	0.113	2.5	39.5	619.5	9704.6	kg
OF Ratio	2.333	2.333	2.333	2.333	2.200
Oxidizer Density	1.949	1.949	1.949	1.949	1.141	g/cc
Fuel Density	0.919	0.919	0.919	0.919	0.582	g/cc
Avg Density	1.640	1.640	1.640	1.640	0.966	g/cc
Average Isp	235	235	235	235	252	seconds
Desired DeltaV	1867.4	1968.3	1968.3	1968.3	3172.7	m/s
Body:Fuel Mass (λ)	0.72	0.63	0.63	0.63	0.197
Payload Ratio	0.047	0.068	0.068	0.068	0.155
Structural Coef	0.419	0.387	0.387	0.387	0.165
Propellant Ratio	0.581	0.613	0.613	0.613	0.835
Mf/Me Ratio	2.249	2.349	2.349	2.349	3.610
Propellant Mass	1.402	22.711	356	5574	52152	kg
Oxidizer Mass	0.981	15.897	249	3901	35854	kg
Fuel Mass	0.421	6.814	107	1672	16297	kg
Oxidizer Volume	503.2	8154.5	127748.1	2001297.1	31423640.1	cc
Fuel Volume	457.3	7410.7	116096.3	1818760.5	28004727.9	cc
Stage Weight	2.411	37.018	580	9085	62425	kg
MT	1.009	14.308	224	3511	10274	kg
Me	1.123	16.832	264	4131	19978	kg
Mf	2.524	39.542	619	9705	72130	kg
Stage Impulse	3230	52338	819926	12844938	128880945	N-s
Cum delta V	1867	3836	5804	7772	10945	m/s

The important thing to notice is that the size of the 3rd stage is quickly too large (close to 620 kg [1370 lbs]). By the time you get to the second stage, it is up to 9705 kg (21000 lbs). The reason is that the performance is too low and weight too high for these motors. For a tiny 113 gram payload, the exo-atmospheric stages are 9705 kg.

SUMMARY

Based on this quick survey of a few commercial rocket motors (yes, it’s a small sample but I think it’s representative), we can see that typical off-the-shelf motors are likely too low in performance and too heavy for orbital launchers. This is not to say that these motors are not of high reliability and capability: they are highly engineered and quality products developed for the commercial market. They are safe and reliable and often reusable. However, these design choices work against them being the lightest possible and what is needed for orbital vehicles.