Ad Astra per Aspera
by Terry Spath
Published 16 Apr 2024

Introduction

Human travel into space is the logical progression of the “Conquest of the Air”. Just as aircraft propulsion advances have preceded aeronautical progress, improvements in rocket propulsion have been key for human exploration of space. The German V-2 rocket was the first vehicle to reach sub-orbital space (defined as Altitude > 100Km.) Its maximum velocity was 3,500 mph.

However, long duration space travel and continuous weightlessness require acceleration from zero velocity on the earth’s surface to 17,500 mph at a minimum. This can establish a spacecraft in a stable Low Earth Orbit (LEO) at 100 miles altitude. Some of the 17,500 mph, “Delta V” can be acquired from the earth’s rotation. Hence, southern launch sites and eastbound launches are advantageous. The Delta V advantage eastbound at Cape Canaveral is 914 mph – i.e. roughly 5% of the required orbital velocity.

Even with the advantage of launching from Cape Canaveral, a propulsionsystem capable of accelerating the spacecraft an additional  16,586 mph (17,500 minus 914 mph) is required. It's said that getting to earth’s orbit is halfway to the planets. There’ a bit of truth in that statement since achieving orbit requires 17,500 mph and escaping earth’s gravity requires acceleration to 25,000 mph or an additional 43% of earth’s orbital velocity.

Mr. Newton in retirement. Young aerospace engineers, including the author, learn to know Mr. Newton well and go to Cambridge University to pay tribute. (Wiki, Wiki)

A common misconception is that orbital weightlessness is caused by absence of earth’s gravitational attraction while in orbit. In a 100-mile-high orbit, the gravitational force is 95% of the value on the earth’s surface. Apparent weightlessness is due to the spacecraft velocity tangent to the earth’s surface which results in a continuous free fall that is equal to the same rate as the earth’s surface is receding. Short periods of weightlessness can be achieved by flying similar trajectories in one of several specially-prepared reduced-gravity aircraft, a.k.a., “Vomit Comet”. Weightlessness in orbit or on the Vomit Comet occurs but an object such as an astronaut’s mass remains the same. Earth’s gravitational attraction is effectively zero at 13 million miles but at that distance gravity from the mass of the sun or other planets will predominate.

We can discover some basic truths about traveling to space with Isaac Newton’s 2nd law, which says Force = mass x acceleration (F = ma). It’s obvious from this simple mathematical relationship that reaching a velocity of 16,586 mph, nearly 3 miles per second, will be easier with a low mass (i.e. the “m”). Airmen are justifiably focused on weight and how it affects aircraft performance. Aircraft engineers routinely utilize a 50% margin of safety (M.S.) when designing and building flight hardware to reduce weight. Many promising aircraft projects have failed to reach their performance goals due to excessive weight growth.

Weight control on space launch systems is even more critical with 20% design M.S.s being the norm. Vast amounts of energy are required to ascend from the earth’s surface, climb out of the earth’s “gravity well” and achieve orbital velocity. The magnitude of the energy required may be appreciated by examining the modified Saturn V that was used to place the Skylab spacecraft in LEO. The Saturn V was the machine that launched Apollo to the moon. For Skylab, the 3rd stage that boosted the Apollo spacecraft to the moon was modified to be an on-orbit manned laboratory and the original 1st and 2nd stages boosted it to LEO.

The Skylab plus Saturn V payload mass fraction of 3% is representative of orbital space launch vehicles. Achieving a positive payload mass fraction has historically been even more difficult for small launch vehicles. This is because material minimum gage sizes prevent minimum weight designs. Also, electronic, pneumatic, and control system weights often do not scale down proportionally to the vehicle size.

The Space Shuttle used cryogenic high performance liquid oxygen (LOX) and liquid hydrogen (LH2) engines for ascent with a payload mass fraction of 5.5%. However, the shuttle had to carry the mass of wings and landing gear. While the mass of the entire vehicle placed in orbit was 240,000 pounds, only about 40,000 pounds was effectively a payload. The remaining 200,000 pounds was the orbiter weight. The SpaceX Falcons and Star Ship recovery systems mostly dispense with airplane like components, but they do have to retain sufficient propellant to manage a powered landing.

Flight vehicle propulsion performance is measured by a parameter called Specific Impulse (I_sp) with units of seconds. An intuitive notion of this measurement can be understood as the duration a propulsion device can produce one pound of thrust while consuming one pound of propellant. The Space Shuttle SRBs operated at an I_sp of 250 seconds. The Saturn V F-1 engine and Space Shuttle Main Engines (SSME) discussed below,delivered I_spof 265 and 440 seconds, respectively. Air-breathing engine performance is sometimes measured with I_sp and values are in the thousands of seconds. However, comparing air-breathing engines to rockets is misleading since the consumption of atmospheric oxygen by air-breathers is not included in the calculation.

Solids

Test firing of the 5 segment SRM for the SLS. 3,600,000 pounds of thrust. Gross weight = 1,600,000 lb; empty weight = 200,000 lb.

Solid Rocket Motors (SRM) have been used for several thousand years beginning with pre-industrial China. Propellant formulations were similar to gunpowder. Prior to the 1950s, small SRMs were widely used as weapons and even for launching rescue ropes to mariners caught in the surf along shorelines. Large SRMs usable for space launch didn’t appear until 1958 when Dr. Dave Altman along with others founded the United Research Corporation (URC) at Menlo Park, CA. URC eventually became part of Pratt & Whitney. URC pioneered the development and testing of large SRMs that eventually led to Solid Rocket Boosters (SRB) used on the Space Shuttle. Even larger SRBs are used on NASA’s Space Launch System (SLS).

Prior to URC’s work, large motors had been problematic because of difficulties encountered when transporting them from their place of manufacture to launch facilities. By developing segmented motor technology, this obstacle was overcome. Segmenting an SRM is more difficult than it sounds. Solid rocket motors use a mixture of a solid oxidizer such as ammonium perchlorate combined with a binder such as Hydroxyl Terminated Poly Butadiene Acrylonitrile (PBAN, a flexible synthetic plastic). Aluminum powder is an energetic fuel and is added although its product of combustion, aluminum oxide at excessive concentrations, can accelerate rocket nozzle erosion. Combustion in an SRM occurs proportionally to the internal exposed surface area. An increase in area such as from a crack or seam can increase the rate of combustion sufficiently to exceed the pressure capability of the rocket motor case resulting in an explosion. URC applied inhibitors at the seams to ameliorate this risk.

In addition to the development of segmented SRMs, URC also developed Liquid Injection Thrust Vector Control (LITVC). A vertical ascending space vehicle requires steering as it flies to follow an ever-changing trajectory to orbit. It starts vertically and then smoothly transitions from vertical to an orbit parallel to earth’s surface. Aerodynamic controls such as fins with hinged control surfaces will only be effective for about the first 10 miles of ascent prior to reaching the vacuum of space. Steering can be accomplished via Thrust Vector Control (TVC). The rocket nozzle and / or complete rocket engine are pivoted around two axes.

The torques and rotation about a single axis.	Typical rocket with TVC mechanical actuation

LITVC is an alternative to mechanical TVC. In LITVC, the rocket jet exhaust is deflected for steering purposes by injecting a liquid into the nozzle exit cone. The liquid is preferably both dense and reactive, such as Liquid Oxygen (LOX). The injected liquid increases mass flow and energy on one side of the nozzle. This process increases thrust in the affected part of the jet producing not only a side force for steering, but additional axial thrust.

My colleague Al Cebriain next to the LITVC nozzle on the AMROC 250,000-pound thrust hybrid rocket ground test rig. The AMROC used LOX for main combustion and some LOX was used for LITVC by injection into the nozzle through the insulated lines connected to the manifold visible to Al’s right.

Liquids

Conceptually, liquid rocket engines are simple devices. Basically, they consist of a pressure vessel into which liquid fuel and oxidizer or a monopropellant such as hydrogen peroxide are pumped and combusted. Similar to solid rockets, liquid rocket combustion chambers contain an exothermic chemical reaction which produces high pressure gas. The high-pressure gas exits this pressure vessel via a convergent - divergent supersonic DeLaval nozzle. The thermal energy of the combustion gases is converted to kinetic energy at the nozzle. If the combustion chamber pressure is approximately 2.5 times the ambient pressure the combustion gases will reach Mach 1 at the nozzle throat and then accelerate further as they expand through the divergent section of the nozzle. Thanks to Mr. Newton, again, and his 3rd law we know that every action produces an equal and opposite reaction. The longitudinal acceleration of the gas molecules, rearward, through the nozzle produces an axial force on the nozzle walls propelling the rocket forward.

Rocketdyne F-1 Engine (Kim McCutcheon)	Sectioned Rocketdyne F-1 Engine MK-10 Turbopump (Tom Fey)

Various propellant combinations and engine configurations have been tried. A spectacular success was the Rocketdyne F-1 engine that burned RP-1 (high grade kerosene) and LOX. On the US moon mission, five F-1s propelled the Saturn V first stage which gave the initial boost to deliver 150 tons to low earth orbit. The Saturn V 2nd and 3rd stages also used bi-propellant liquid engines that burned cryogenic liquid hydrogen and LOX. The 3rd stage propelled the Command Module, Service Module, and Lunar Module (LM) to lunar orbit where Neal Armstrong and Buzz Aldrin piloted the LM to human’s first landing on the surface of an extraterrestrial body on 20 July 1969.

Bi-propellant rocket engine technology was initially pursued by the Germans during WWII with the best-known example being the V-2 that was used to attack Britain. Many German engineers came to the US after the war under Operation Paperclip. Perhaps, the most prominent was Dr. Wernher von Braun who was very involved in the U.S. moon flights. Not surprisingly, the F-1 technology to some degree paralleled the German V-2 engine. Both combusted hydrocarbon fuel with LOX and used turbine-driven fuel and oxidizer pumps.

A gas generator is similar to the combustion section of an aircraft turbine engine except air feed from a compressor is replaced by an injector that mixes a liquid oxidizer and fuel. The F-1 used RP-1 and LOX. The combustion gas output from the gas generator is routed through a turbine that is mounted on the same shaft as the fuel and oxidizer pumps.

The F-1 nominal main combustion chamber pressure was1,125 psi and the turbopumps had to overcome that pressure to feed the propellants through the injector into the combustion chamber. Sufficient “delta P” across the injector was assured by the fuel pump output of 1,856 psi. Net positive suction head pressure to prevent pump cavitation was accomplished by pressurizing the RP-1 and LOX tanks to approximately 30 psi.

The gas generator delivered 171 pounds per second of 1,450° F combustion gas consisting mostly of CO2 and water, which are the products of combustion when burning a hydrocarbon fuel with oxygen. The gas pressure was 905 psi as it entered the turbine, and this allows it to deliver 52,900 horsepower to the pumps. The LOX pump, RP-1 pump and turbine were stacked in that order on a common shaft driven by the turbine that spins at 5,488 rpm. The gas generator initially starts with the previously noted propellant tank ullage pressure of approximately 30 psi and as the pump delivery pressure increases, the gas generator propellant inlet pressures also increase since they are tapped off of the pumps’ outlets. Incidentally the General Electric J79 turbojet used in the McDonnel F-4 has a mass flow rate of 175 pounds per second, nearly equal to the F-1 turbopump gas generator. This author has stood near a J79 at full thrust and is impressed to consider that this amount of conspicuous energy was used just to pump propellant into the F-1.

More detail on operation of the Saturn V 1st stage was detailed in Mr. Tom Fey’s AEHS article; One Second in the Life of the Rocketdyne F-1 Rocket Engine

The U.S. Space Shuttle used a combination of solid rockets and bi-propellant liquid engines firing in parallel to boost the 240,000-pound Orbiter to LEO. The Shuttle Gross Lift-Off Weight (GLOW) is 4,400,000 pounds. Because the Shuttle was designed to be partially reusable, heavy components such as wings and landing gear were used for return to earth. These masses also had to be accelerated to orbital velocity in addition to a payload. To achieve a reasonable payload mass fraction, the Space Shuttle Main Engines (SSME) engines had to be very efficient. Efficiency, as measured by I_sp,improves by increasing the ratio of chamber pressure to the external ambient pressure. This can be noted by the fact that the SSME combustion chamber pressure was 3,006 psi compared to the 1,125 of the Saturn V F-1s.

The AEHS has a comprehensive account of the development effort that led to operational SSME. This author worked at Rockwell while this effort was underway and remembers the numerous test stand failures detailed in the articles which were undoubtedly a result of the large increase in rocket technology state-of-the-art that was underway.

As noted above, the SSMEs were some of the highest performing liquid rocket engines ever flown and delivered an I_sp of 440 seconds. By comparison, the Saturn V F-1 engine’s I_sp was 264 seconds.

The SSMEs were designed to be reusable and after extensive inspection and refurbishment after each of the Shuttle flights they were indeed re-flown. Previous rocket engines such as the F-1 were dropped into the Atlantic during each Saturn flight. Now the NASA Space Launch System (SLS) is being flown with four one-time-use derivatives of the SSMEs designated as RS-25s. Redesign of these derivatives includes techniques to lower their cost since they will be disposed of during each flight. Recent reports about the RS-25 engines say their cost will be roughly $100 million per engine.

Six RS-25 engines for the SLS

Hybrids

Combining elements of liquid and solid propulsion, the most common hybrid rocket motors use a liquid oxidizer and solid fuel. Common liquid oxidizers are LOX, hydrogen peroxide, and nitrous oxide (N2O). N2O is advantageously used since it is self-pressuring to 750 psi at room temperature. This allows a simple feed system consisting of an oxidizer tank that can withstand that pressure with few other components, mainly, a valve that is opened to admit the N2O into the combustion chamber after ignition. LOX and hydrogen peroxide oxidizers require a separate pressurizing gas to overcome the combustion chamber pressure. High-pressure helium gas that is regulated down to the tank pressure is one commonly used example.

Simple hybrid vehicles that depend solely on N2O’s self-pressurization characteristic do experience a thrust decrease as the N2O tank is emptied. Prior to operation the N2O exists in equilibrium with liquid and gas fractions determined by their temperature. As N2O is depleted and the tank ullage expands during operation the remaining N2O cools,resulting in a decrease in its vapor pressure. This may be an advantage since rocket mass is decreasing during flight and a decreasing thrust profile may hold the acceleration constant. Certain payloads may be damaged by high acceleration forces. Referring again to Mr. Newton, F = ma, and with a decreasing m and decreasing F, a can stay constant.

N2O is generally recognized as a relatively safe chemical, and it is used as anesthesia by dentists and a pressuring agent in whipping cream cans. Undoubtedly, N2O is safer than other types of rocket propellant such as hydrogen peroxide or hydrazine, both of which can cause life threatening trauma. However, N2O exothermically decomposes, meaning that net positive energy is delivered when sufficient activation energy, such as a hot surface, is encountered.

In oral surgery, blood vessel heat cauterization is hazardous when using N2O as an anesthetic since a heat source is present in the mouth along with N2O. Fast acting valves in a N2O system may generate sufficient heat from poppet friction to initiate N2O decomposition. This author participated in the investigation of a fatal N2O decomposition event. One example in which the participants were lucky to survive an accident unscathed occurred as follows: An N2O delivery was underway from a tanker truck to an industrial facility. Pumping was required from the truck since the truck tank and storage tank were at the same temperature. Since several thousand gallons of N2O needed to be transferred, the truck driver walked to the front of the truck to take a smoke break. A pump bearing overheated, and a decomposition explosion resulted. The tank was destroyed but the driver, ironically, was saved by smoking.

Nonetheless, hybrid rocket vehicles have the potential to be considerably safer and simpler than solids or bi-propellant liquid powered craft. The fuel for an SRM is generally classified as a 1.3 non-explosive compound. However, this fuel will energetically combust and is nearly impossible to extinguish if accidentally ignited. Also, SRMs are almost impossible to design with throttle ability. Bi-propellant liquid rockets may be throttleable but suffer from extensive mechanical complexity due to pumps, valves, and associated plumbing. Hybrids generally exhibit low explosive danger until the introduction of the liquid oxidizer into the combustion chamber and are simpler than comparable bi-propellant liquid rockets. They also can be throttled by controlling the oxidizer feed rate.

Hybrid fuels typically are some class of pure hydrocarbon solid, although energetic components such as metallic powder may be added to increase the performance. Early work on hybrid fuels was conducted at Cal Tech and, interestingly, wood was even investigated as a potential fuel material. Petroleum based solids such as Nylon and Hydroxyl Terminated Poly Butadiene (HTPB) are commonly used for hybrid fuels. HTPB is also used as a binder in solid motors but its performance in hybrids is lower. The internal combustion process in a HTPB hybrid requires the oxidizer to penetrate a carbon rich char layer that forms during combustion on the fuel surface. In a solid, the oxidizer is in intimate contact with the fuel surface by virtue of being a component of the fuel / oxidizer solid fuel grain.

Paraffin fuel hybrids were investigated by Space Propulsion Group (SPG), a spinoff company from a Stanford University Aerospace PhD thesis. This author participated in developmental tests for SPG over a period of 15 years. Paraffin hybrids appear to have better performance than HTPB and nylon. A char layer does not form on the fuel surface since the fuel melts and the molten fuel forms droplets that are combusted while entrained in the combustion gas flow path. SPG research rockets in static tests demonstrated I_spin excess of 300 seconds.

SPG closed its doors in 2019 but not before making good progress on a Paraffin / Mixed Oxides of Nitrogen (MON) hybrid to propel a rocket from the surface of Mars. The NASA Perseverance Mars rover is currently traversing the surface and packaging and sealing small vials of rock and Mars regolith. At a future date, another robotic mission will be landed on Mars. The samples produced by Perseverance will be collected, loaded onto a Mars Ascent Vehicle (MAV) and flown to rendezvous with an orbital spacecraft that will return the samples to Earth. The rocket for the MAV must be capable of surviving on the Martian surface where temperatures can be lower than minus 100°F, a temperature capable of permanently rendering most rocket propellants unusable. SPG conducted extensive tests with a specially developed paraffin that would survive this environment and then burn successfully with MON to propel the MAV to orbit.

Letara, a Japanese aerospace propulsion company also grew out of a PhD thesis, at the University of Hokkaido. Their hybrid is a HDPE / N2O rocket that will serve as “kick motors” for changing satellite trajectories and orbits. One core technology which needed to be developed to enable hybrids to perform such maneuvers is a low-power and energy efficient method of achieving multiple re-ignitions. Existing igniters would be consumed during operation and require power sources which exceed the capability of the typical small spacecraft.

The flights of SpaceX Falcon 9 and super large bulk carrier vehicles such as the SpaceX Star Ship presage reaching LEO for significantly lower cost per pound. These large boosters will contain tens if not hundreds of small payloads. Falcon 9 flights are already launching 23 Starlink satellites per mission. Future small satellite missions will be able to reach high energy trajectories utilizing Letara tugs to reach the moon, Mars, and beyond.

Letara’s upper stage motors offer reduced risk for use in a large boost vehicle due to their low toxicity oxidizer and non-hazardous HDPE fuel. Similar to maritime freighters carrying a large number of small items with hazards such as lithium batteries, an entire Star Ship mission could be jeopardized by a “Black Swan” event onboard one small payload. An inadvertent SRM ignition or leak of a toxic reactive propellant such as hydrazine could destroy the entire vehicle.

This author was alive and watched John Glenn launch into orbit on a TV in a 6th grade classroom. I’m optimistic that humans will be on their way to a multi-planetary existence in my lifetime.

References

Pioneers in Propulsion — History of CSD Pratt & Whitney’s Solid Rocket Company
Turbopump Systems for Liquid Rocket Engines (Cleveland, Ohio: NASA Lewis Research Center, August 1974)

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