Low Earth Orbit Atmospheric Scoops

The rocket equation is often described as tyrannical. Low exhaust velocity chemical rockets that are used today need payloads to be mounted on towers of propellant to reach extraterrestrial destinations. 

Re-fueling a rocket in orbit has been floated as a solution to drastically reduce rocket sizes. The best source of propellant for this purpose is much closer than is generally thought... 

Orbital refuelling

Orbital propellant depot

Orbital refueling is a concept where a spaceship's propellant tanks are refilled after it achieves orbit. Without refuelling, a spaceship must carry the propellant it needs to reach its extraterrestrial destination up from the ground. With refuelling, it can launch empty or even reuse a propellant tank once in orbit. 

The propellant can be launched separately or manufactured from extra-terrestrial sources such as the Moon or an asteroid. In this post, we will be looking at a way to simply scoop up propellant from the atmosphere.

Beating the rocket equation

A quick calculation reveals how advantageous orbital refuelling is.

SpaceX's Mars mission architecture relies on orbital refueling for good reason.

Imagine a 10 ton spaceship that we want to send to Mars. It needs 9.5km/s of deltaV to enter orbit, and another 5.6km/s to go to Mars. Let us imagine that it uses engines with an average specific impulse (Isp) of 350 seconds, which corresponds to kerosene-liquid oxygen rockets. We can also write this as an exhaust velocity of 350*9.81: 3433m/s. 

The rocket equation is as follows:

• DeltaV = Exhaust velocity * ln(Mass Ratio)

DeltaV, in m/s, is the velocity a rocket can reach if it expends all of its propellant reserves. Exhaust velocity, in m/s, is how fast a spaceship's engines expel propellant. Mass ratio is the fully fuelled or 'wet' mass divided by the empty or 'dry' mass. 

We can re-write that equation to work out the mass ratio required to obtain a certain deltaV.

• Mass ratio = e^(DeltaV/Exhaust Velocity)

e is the exponent function, roughly equal to 2.718. If we want to reach orbit, we need 9500m/s of deltaV. Using an average exhaust velocity of 3433m/s, we calculate that the mass ratio required is e^(9500/3433): 15.9

This means that the wet mass is 15.9 times greater than the dry mass on the launchpad.

Going from orbit to Mars requires a mass ratio equal to e^(5600/3433): 5.1

DeltaV requirements for travel between Earth, Moon and Mars.

The total mass ratio on the ground is the product of both mass ratios. 15.9*5.1: 81. If our spaceship is 10 tons, then the rocket on the launchpad weighs at least 810 tons. A realistic spaceship with propellant tanks, structural support and engines will cut into the available payload mass. We will end up with a towering rocket that is mostly propellant, with a tiny cargo bay at its tip. 

What if we use orbital refuelling? 

How reusability and orbital refueling can make a Moon mission drastically smaller 

The mass ratio for launch remains the same. However, we do not need to carry up propellant from the ground to go to Mars with it - instead, we refill the same tanks with more propellant in orbit. 

The mass ratio on the ground is therefore 15.9. The rocket will mass only 159 tons on the launchpad. This is a rocket more than five times smaller than the previous example. 

Using the same calculations with the Moon as destination (3300m/s) still means the rocket is e^(3300/3343): 2.61 times smaller than without refuelling. 

Current challenges and solutions

ITS and its tanker docking in LEO to transfer propellants.

The reason orbital refuelling has not been used to drastically reduce the size of rockets on the launchpad is that currently, the only way of delivering propellant to a spaceship in orbit is with another launch. This negates any direct advantage refuelling provides. 

Building a lunar base.

The other option discussed as a source of propellants is in-situ resource utilization. This involves setting up an industrial base on the Moon, an asteroid or elsewhere to produce propellant from locally available materials. The products are shipped back to Low Earth Orbit for refuelling spaceships. It makes sense from a deltaV perspective: it is easier to send propellant back from Mars (5.6km/s) than it is to lift it up from Earth (9.5km/s)!

However, it would require a massive investment in extraterrestrial infrastructure that would not help anyone for several decades, requiring automated technology that can work independently or a permanent manned settlement. 

China Aerospace Science and Technology Corp.'s 30km/s Hall-effect ion thruster.

NASA has been working on alternatives to orbital refuelling for high-deltaV missions. The best solution so far is the use of high exhaust velocity solar-electric propulsion.

Dawn approaching Ceres.

As noted above, mass ratio depends on exhaust velocity. A higher exhaust velocity drastically reduces mass ratio. Solar electric rockets such as Dawn's ion engine have an Isp of 3100 seconds. Other electric designs are capable of even greater performance, powered possibly by nuclear reactors. 

High Isp rockets could reduce the mass ratio required for a mission to mars down to e^(5600/30411):1.2 or less. It would be more than four times smaller than a chemically-fuelled rocket.

Atmospheric Gas Scooping

This concept consists simply of running a gas scoop through the upper atmosphere and collecting the atmospheric gasses to be used as propellant. Some of the propellant is consumed by the scoop itself, the rest of made available for orbital refuelling of other craft.

The PROFAC gas scoop operated at 120km altitude and was powered by a 10MW reactor (4) and had some rocket fuel on-board to burn with collected oxygen to produce thrust in case the reactor failed. 

We have to work out some orbital mechanics to find out how much gas the scoop-equipped spacecraft gets to keep. In essence, it is ramming a scoop through immobile air and accelerating it to orbital velocity. The scoop loses momentum - it must expel some of that gas as propellant through a rocket to recover that momentum. 

The momentum a scoop loses by collecting 1 kilogram of air at low altitude is 1*7800: 7800 Newton-seconds of momentum. The momentum gained by using a rocket is equal to exhaust velocity*propellant flow. The exhaust velocity*propellant flow product must be greater than 7800 Newton-seconds. 

If the exhaust velocity is lower than orbital velocity, the scoop will need to consume more propellant than it could ever collect to stay in space.
If the exhaust velocity is equal to orbital velocity, the scoop has to use all of the gasses it collects as propellant and retains none. This is the case for solid-core nuclear thermal rockets.
If the exhaust velocity is higher than orbital velocity, then the scoop will be able to retain some of the gasses in reserve and produce a net gain in propellant collected. 

The ratio between propellant retained and propellant consumed is simple. A rocket like Dawn's ion engine will produce 1*30411: 30411 Newton-seconds of momentum per kilogram of propellant. It can therefore retain up to (30411-7800)/30411: 0.743 or 74.3% of the gasses it collects.

The gas retention percentage is simply calculated as:

• Retention % = [(Exhaust Velocity - Orbital Velocity)/Exhaust Velocity]* 100

We can quickly work out that propulsion systems with an exhaust velocity just above orbital velocity will have a very hard time collecting propellant. An ideal engine has an exhaust velocity several times the orbital velocity. However, as we will now see, getting enough power to these engines is problematic.

Powering the scoops against drag

Running an electric rocket requires a good amount of power. In space, this is hard to come by.

The power requirements of a gas scoop are determined by the thrust it needs to produce, which depends in turn on the drag it experiences. Aerodynamics play no role at orbital velocities, drag is only a function of cross-section and atmospheric density. This equation will give you an estimate within +/-10% of reality. 

• Drag: 0.5 * Orbital velocity^2 * Cross-section area * Gas density 

Drag will be measured in Newtons.
Orbital velocity will be in m/s. 

Gas density chart up to 300km altitude from here. 

Cross-section is the frontal surface area (m^2) of the gas scoop ship that intersects with the gasses it is traversing. Some of it will be the scoop opening, into which gas is funnelled, some of it will be closed surfaces that gas bounced off of. A gas scoop ship will want its funnel to cover the entire frontal area to minimize wasteful drag (drag not contributing towards collecting gasses). Therefore, the cross-section area is the funnel opening area.

Gas density is determined by altitude and noted in kg/m^3. If the scoop operates at a single altitude, we use a single value. If it changes altitude, we use an average value based on the time spent at each altitude. At very high altitudes, some sources will note a value in particles per cubic meter. This has to be converted into kg/m^3. 

Particle counts for different elements per cubic meter for altitudes up to 1000km. 

Notice how mass has nothing to do with drag or the thrust requirements. Gas scoops can be as massive as they want so long as they are able to handle drag. Thrust requirements will not change as it collects gas either, nor will they be lower when the scoop's tanks are empty. 

So how much power do we need?

We first must estimate the drag generated and give the scoop ship enough power to produce sufficient thrust to counter the drag, plus a safety margin.

The ISS reboosts its orbit using a Zvezda module's engines. It could be performed by electric rockets instead. 

Here is a table of drag values listing the drag force per square meter of cross-section area at different altitudes, for 7800m/s orbital velocity, using data from this NASA website:

We can see that above 200km altitude, the drag force is in millinewtons, descending down to tenths of micronewtons per square meter at the edge of the atmosphere. A more accurate reading would take into account the small differences in orbital velocity as the altitude changes. It is 7800m/s at 150km altitude, but reduces slightly to 7350m/s at 1000km altitude.

The equation for engine power is the following:

• Engine Power: Thrust * Exhaust Velocity /2

If we combine it with the drag equation, we obtain this:

• Engine Power: Drag force per m^2 * Cross-section * Exhaust Velocity /2

Engine power is in Watts. Thrust and Drag force per m^2 are in Newtons. The Cross-section area is in m^2 and the Exhaust velocity is measured in m/s.

Example:
10m wide funnel at 200km
Cross-section area: (10/2)^2 * 3.14 = 78.5m^2
Drag force per m^2 at 200km: 9.51 * 10^-3 Newtons
Exhaust velocity: 30411m/s
Engine power: 11.35kW

100m wide funnel at 400km
Cross-section area: (100/2)^2 * 3.14 = 7850m^2
Drag force per m^2 at 400km: 1.03 * 10^-4 Newtons
Exhaust velocity: 30411m/s
Engine power: 12.3kW

The engine power here is the effective power output going out of the nozzle. Rocket engines are not 100% efficient, and there are further losses in the systems that generate, transport and convert electricity going into the engine. These might double the actual power consumption. 

There are many options for producing the required energy to power the engines, but only a few are practical or achievable with the technologies available in the near future. 

Apollo's fuel cells. 

Chemical energy, such as in a fuel cell, is not a viable option. The system masses too much and requires constant refuelling. Radio-isotopes are compact and reliable, but their output is very, very low at only about 2 to 5kW per ton. 

The two options remaining are solar panels and nuclear reactors. 

Solar panels are cheap and lightweight, able to produce over 580W/m^2 at peak efficiency. However, they can become very large if we need kilowatts of power. They must be packed behind the main funnel or they would increase the cross-section area. A 10kW array of solar panels might be divided into eight segments 2m wide and 1.07m long. Thin-film solar panels might reduce mass down to a few dozen kilograms. 

Nuclear reactors are very powerful and are unlikely to need much mass or volume to produce the output required. Experiments have been conducted and prototypes have been flown of space-rated nuclear reactors, but historical and political reasons have prevented their widespread use. We can send a 100kW reactor today into space massing only 512kg.

Gas collection and processing

A turbomolecular pump

A funnel scoops up gasses by ramming through the air and compressing them into a tube. In that tube, pressures rise until it reaches stagnation pressure. This is the pressure when the gasses have been completely stopped relative to the scoop. Because the scoop is travelling so much faster than the particle velocity inside the gasses being collected, it can act as a giant piston like in a turbomolecular pump. The scoop naturally compresses gasses while they cannot escape back out of the opening. 

The first step is to move the pressurized gas into an empty gas tank. It can be vented without any pumping equipment and only a valve to maintain the flow going one way. As long as the collection tanks are kept at a lower pressure, no active pumping is needed. 

To keep collection tanks at a lower pressure, the gasses must be cooled until they become liquid, then they are transferred to an insulated long-term storage tank. This is the second step.

Air intake and compression are tasks performed by the scoop during its operation.

Cooling the collected air until it liquefies allows for its fractional distillation into different gasses. The gasses have different boiling points and separate into different fluids as the temperature drops. The relative proportions of each gas this fourth step produces is determined by the atmospheric composition at the collection altitude. 

Number of N2 particles per cubic centimeter at altitudes 100 to 1000km

Number of O2 particles per cubic centimeter at altitudes 100 to 1000km

Number of atomic nitrogen N particles per cubic centimeter at altitudes 100 to 1000km

Number of atomic oxygen O particles per cubic centimeter at altitudes 100 to 1000km

Number of atomic hydrogen H particles per cubic centimeter at altitudes 100 to 1000km

As noted before, we must convert particle counts into kg/m^3 densities.

For example, at 200km altitude, we note that we get 3.8*10^9 N2 particles, 3.25*10^8 O2 particles, 4.39*10^9 O particles, 9.62*10^6 N particles and 2.15*10^5 H particles per cubic centimetre. We multiply the values by a million to get the cubic meter values, then convert the particle numbers into masses by multiplying by the element's molar mass in g/mol then dividing by Avogadros' constant (6.022 * 10^23). 

We can work out the following figures:

N2 particle mass: 28.0134 / 6.022 * 10^23 = 4.65 * 10^-23 grams 

N2 mass per m^3 at 200km: 1.78 * 10^-16 kg

O2 particle mass: 31.9988 / 6.022 * 10^23 = 5.313 * 10^-23 grams 

O2 mass per m^3 at 200km: 1.72 * 10^-17 kg

O particle mass: 15.9994 / 6.022 * 10^23 = 2.65 * 10^-23 grams 
O mass per m^3 at 200km: 1.16 * 10^-16 kg

N particle mass: 14.0067 / 6.022 * 10^23 = 2.325 * 10^-23 grams 

N mass per m^3 at 200km: 2.23 * 10^-19 kg

H particle mass: 1.008 / 6.022 * 10^23 = 1.67 * 10^-24 grams 

H mass per m^3 at 200km: 3.59 * 10^-21 kg

What can we understand from these figures? 

The atmospheric composition at 200km is dominated by nitrogen and monoatomic oxygen. Density when all gasses are considered is 3.127*10^-16kg/m^3... nitrogen therefore represents 56.9% of the mass of air collected at this altitude, with monoatomic oxygen second at 37%. The relative proportions of the gasses allows us to estimate the average heat capacity; it should be about 1.2kJ/kg/K.

Stirling engine animation

It is important as it tells us how much energy will be needed to reduce the temperature of the gasses down to their liquefaction values. A Stirling-engine cryocooler is suited to the task, with a pump efficiency of roughly 40%.

Here is a table of the temperatures, energies and power consumption levels involved in liquefying air collected at 7800m/s for altitudes between 100 and 1000km:

Note that the power ratings are in milliWatts per square meter of scoop cross-section area. It is the electrical consumption for a heat pump cooling down the gasses from the collection temperature to 77K, where nitrogen liquefies.

For 200km, we read that 19.48W per square meter is required. A 10m wide scoop would need to devote only 1.5kW to cryocooling at this altitude. 

Gas scoop design example

We will attempt to work out a ballpark estimate of the mass, size and performance of a conventional gas scoop that uses currently available technology. 

This simple design operates at one altitude and continuously uses its engines counter the drag forces of ramming through the air.

We will stay within the limits of a payload that could be launched by a single Falcon 9 FT rocket to LEO, so 22.8 tons to 200km altitude.

Let's start with the engine. We want it to use the gasses it collects as propellant. This excludes current electric rockets propelled by xenon or argon. Instead, we need to find examples of designs that are happy to run on nitrogen. Oxygen is viciously corrosive when hot, so should be avoided as a propellant despite its relative abundance. 

Diagram of the electrodeless RF plasma thruster.

Looking at this engine list, we find the electrodeless RF plasma thruster. It has the ability to accelerate any propellant to extreme velocities at high efficiency, without the temperature or endurance limits of other electric rockets. 23km/s exhaust velocity with nitrogen as propellant can be achieved by a plasma with an electron temperature as low as 5eV, and an ion temperature of 25eV. 

The VASIMR rocket engine by AdAstra.

The most similar engine we have mass and efficiency data on is the VASIMR, with 580W/kg in low gear and 60% overall efficiency. We can work out that each kW of electrical power supplied to the RF plasma thruster produces 0.052N of thrust out of 1.72kg of engine.

As mentioned before, solar electric power is the most realistic option. Supplying 1kW of electricity using thin-film technology requires 2.02m^2 of solar panels and about a kilogram of electrical equipment after including 15% losses during electrical conversion and transport to the engines.   

Putting all these figures together gives us a figure of 34kg and 2.02m^2 of power and propulsion per Newton of thrust. 

At 200km altitude, the orbital velocity is 7790m/s. Air density is 3.127*10^-10 kg/m^3. The drag force per square meter of scoop cross-section area is 9.487*10^-3 N/m^2. 

Using the power and propulsion figures, we determine that each square meter of scoop area requires roughly 0.326kg of equipment to counter drag forces. 

A drag force in the range of milliNewtons allows for the use of very lightweight scoop materials. The closest analogy would be a hemispherical balloon holding up to a pressure differential, with structural support lines transmitting load or tension forces to a payload. Stratospheric balloons survive 200Pa pressure differentials and carry several tons using membranes massing 55 grams per square meter.

We might be able to say that despite the hundred-thousand-fold decrease the in the forces involved we have decided to include massive redundancies against micrometeorite impacts  and friction ablation so the scoop will mass 0.05kg/m^2, similar to a stratospheric balloon.  

Processing the gasses requires some more equipment. 2.44 micrograms per square meter per second. are collected at this altitude and velocity. 94% of that is N2 and O, reducing the collected mass to 2.29 * 10^-9 kg/m^2/s. It requires a heat pump and gas handling system able to handle 18W of power. SABRE reaction engines claims to be able to handle 400MW of cooling with a one ton device, but we will use a much more conservative 400W/kg. 18W of cooling would require 0.045kg of equipment.

The total so far is 0.426kg/m^2.

We want our gas scoop to collect gasses for one year, then offload them. We therefore need storage tanks able to handle liquid oxygen and nitrogen for one year. They store at similar temperatures, so we will use 5% of the propellant mass as tank mass, something based on this NASA report for the Jupiter launch vehicle. It is based on Aluminium-Lithium materials, but even more lightweight carbon fibre is possible. 

World's largest carbon-fibre tank for holding liquid oxygen.

Liquid oxygen has a density of 1141kg/m^3. Liquid nitrogen has a density of 808kg/m^3. Over a year, a gas scoop orbiting at 200km altitude produces 4.37*10^-2kg of liquid nitrogen and 2.84*10^-2kg of liquid oxygen, per square meter of scoop area. This requires a combined storage volume of 7.89 * 10^-5 m^3, massing a negligible 3.65 grams. 

Due to the 23000 to 7800 ratio between the engine's exhaust velocity and orbital velocity, 33.9% of the gasses collected must be expelled as propellant. This will reduce the amount of nitrogen collected from 56.9% to 22.98% of the total mass of gasses collected. This reduces the nitrogen storage volume required by a factor 2.47 and the yearly tank mass to 4.67 * 10^-5 m^3, and the mass to 2.3 grams.

With all functions accounted for, the scoop will require 0.428kg of equipment per m^2 of cross section area. If we double this figure to account for everything including attitude thrusters, docking structures, thrust frames, communications, propellant handling, safety margins and so on, it is still about 0.857kg/m^2. 

Now we can derive the maximum size of this scoop and its yearly performance.

22800kg / 0.857kg = 26604m^2

A single Falcon 9 FT can put a scoop with a cross-section area of 26604m^2 into orbit. It would represent a disk 184 meters wide, although multiple smaller scoops would be more sensible. They would be equipped with thrusters producing a total 252 newtons of thrust and use 509m^2 of solar panels. 

These scoops would collect 2046 tons of gas, of which 471 tons of liquid nitrogen and 757 tons of liquid oxygen is retained, for a total of 1228 tons.

What's the point of all this?

Remember the mass ratio equations at the beginning of this post? A mass ratio of 15.9 to reach orbit can be understood as each kilogram of propellant in orbit being worth 15.9 times more than the same kilogram on the ground.


1128 tons in orbit would be worth 19525 tons on the ground. These thousands of tons could be saved without having to capture an asteroid, mine the moon or invent a radically new launch system. They alone would make the entire concept of atmospheric gas collection worthwhile.

Understandably, the gasses collected cannot be simply loaded into rocket propellant tanks and expected to burn. Liquid nitrogen is inert and cannot be 'burned'. Liquid oxygen needs a fuel, which can be liquid oxygen, methane or kerosene lifted up from the ground or already on-board the spaceship expecting to be refuelled. 

In a typical LH2/LOX rocket engine, six kilograms of oxygen are consumed for every kilogram of hydrogen. Looking at it another way, for each kilogram of liquid oxygen a spaceship receives through orbital refuelling, it needs to carry along 0.167kg of liquid hydrogen from the ground or separate launches.

A 10 ton dry mass spaceship being sent from low earth orbit to geostationary orbit would require a deltaV of 3931m/s. If it uses a LH2/LOX 450s Isp rocket engine, it would require 14.3 tons of propellant. It would need to bring up 2.39 tons of liquid hydrogen from the ground, but would save on 11.96 tons of liquid oxygen propellant. 
The lifter on the launchpad would be 49% lighter. 
A single Falcon 9 FT-launched scoop could provide enough liquid oxygen to push 633 tons of payload from LEO to geostationary orbit per year.
For comparison, a 353s Isp kerosene/liquid oxygen rocket with an oxidizer/fuel ratio of 2.56 would reduce lifter mass on the launchpad by 52% and is enough for 498 tons to GEO.

There are further savings if we consider that that launcher would need smaller propellant tanks, structures and fewer engines when it is nearly two times smaller. LEO to GEO is a common requirement for satellite launches. These savings might allow for much more realizable Lunar or Martian missions. 

Other secondary benefits exist.

The name 'Orbital Transfer Vehicle' or OTV is given to the rocket stage that moves a payload from low earth orbit to geostationary or lunar orbits. EDS stands for 'Earth Departure Stage' for payloads being sent to Mars or elsewhere. An electric or nuclear-thermal OTV or EDS can use all of the gasses collected (1228 tons) and can essentially launch with empty tanks, meaning that the full advantages of orbital refuelling can be reaped. 

The availability of hundreds to thousands of tons of propellant waiting in orbit will incentivise the development of non-chemical propulsion technologies. 

Alternative designs and improvements

The two heaviest components of an atmospheric scoop are power and propulsion equipment. 

Here are some solutions to reducing their mass requirements so that even bigger scoops which can collect even more gas per year can be launched on the same launcher:

-Use all the nitrogen:

If atmospheric gas scooping is going to provide oxidizer to spaceships in low orbit, then only liquid oxygen is of interest. All the liquid nitrogen can be converted into propellant on-board the scoop. At 200km, this allows for up to 57% of the gasses collected to be consumed, meaning Isp can be as low as 18km/s and thrust per kg will increase by a third compared to our example. 

Furthermore, if the nitrogen is being consumed as it is collected, then there is no need to liquefy it and less cryocooling equipment will be needed. 

-Use better power density and space-grade components:

In our example, the power density of the electric rocket engine was 580W/kg, cryocoolers were 400W/kg and we doubled the mass per square meter in the end. Specialized technology using space-grade materials could greatly reduce the mass per square meter. 

-Trawler design

Instead of fish, it would be collecting gasses from a lower altitude.

To fulfill all the energy needs on-board our scoop design, we needed to equip it with 509m^2 of solar panels. Keeping these panels and other equipment hidden behind the scoop's funnel is challenging, as it elongates the design and increases structural requirements to prevent bending and flexing. 

If the main body of the design and its scoop funnel were physically separated, the overall design could be lighter. An arrangement where the funnel is dragged at a lower altitude by tethers and pumps connecting it to the main body is called a 'trawler', as it resembles boats dragging along nets under the sea. For one, solar panels can be arranged any way desired. Second, a scoop traversing lower altitude air can collect more gas per square meter, allowing for a smaller cross-section for the same performance and lower mass per square meter. Third, the scoop ship does not have to consume a lot of propellant to lift itself out of a lower altitude if it wants to stop scooping: it can simply winch up the funnel. 

The disadvantages are greater complexity and the need to pump gasses from where they are collected to where they are processed and stored.

-Diver design

A conventional or trawler gas scoop design needs engines to counter-act drag with an equal amount of thrust at all times. Solar panels must provide sufficient energy and store it to keep propulsion running around the night side of Earth. 

There is a design that allows a scoop ship to get away with much smaller engines and power sources. A 'diver' operates at an elliptical orbit, dipping into the atmosphere only at its lowest point. At that point, it rapidly scoops up gas and rams its way up and out with its momentum. For the rest of the orbit, it fires up its small engine to recover the lost momentum over a long period of time. Much smaller solar panels would be needed to feed propulsion and gas processing equipment. If the diver goes into the lower atmosphere with excess velocity, like an aerocapture maneuver, it could use a comparatively tiny scoop to collect a lot of gas, so overall the diver scoop will be much smaller and lightweight compared to other designs.

If it needs more time, the scoop can perform one diver every several orbits. 

The downside is that ramming through the lower atmosphere at higher than orbital velocity involves significant heating which lowers the ship's useful life. Collecting gasses quickly means that some intermediary store must be available before they are processed and liquefied: holding large volumes of hot gas would require voluminous tanks that add drag and weight to the scoop ship.  

-Electrodynamic tethers

Scoop ships spend a lot of time in space, have a lot of electrical power available and only need tiny thrust output to maintain their altitude. Electrodynamic tethers are perfectly suited to atmospheric gas scooping operations. They push against the Earth's magnetic field to provide a form of propulsion that does not require any on-board propellant. Electrodynamics tethers can supplement or replace electric rocket engines.

-Electric scoop

Between 60 and 1000km altitude, Earth's atmosphere stops being a continuous medium and becomes a loose plasma dominated by charged particles. This is called the ionosphere. For example, monoatomic oxygen is a negatively charged particle that represents a large fraction of gasses at 200k altitude and above. 

These ions can be collected by an electromagnetic/electrostatic scoop. The electromagnetic section is composed of large magnetic fields that direct ions towards the ship from a large volume of space. It would be similar in operation to a Bussard Ramjet. The electrostatic section is composed of extremely thin rings with a charged interior. Ions of the unwanted charge are deflected, ions of the desired charge are pulled to the center, where a physical scoop collects the gas. The main advantage of this design is that it allows a tiny scoop and a set of magnets to do the work of a much larger funnel.  

An interstellar voyage by the Bussard Ramjet

An electric scoop would be the only practical way to collect significant amounts of hydrogen from altitudes above 400km. Scoop cross-sections several kilometers in diameter would have the same performance as a scoop a hundred meters wide at lower altitudes. With an electric scoop, such diameters are possible without scaling the mass to cross-section area.