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Design and Application of Synergetic Air Breathing Rocket

Joseph Pointer1 University of Colorado, Boulder, CO, 80309

May 11, 2017

The following report provides an overview of the design, operation, performance, feasibility, and application of a synergetic air breathing rocket (SABRE). The system is discussed in detail to provide understanding of the benefits the technology has to offer to the commercial transport and space launch industries alike. A view of the technological development is provided in terms of the efforts currently underway to develop and prove the feasibility of this system. The SABRE is shown to be a promising technology, but has several design obstacles to overcome before it can enter the market as a viable competitor to current propulsion systems.

I. Background HE concept of a single stage to orbit (SSTO) vehicle is an attractive idea to competitors in the Tcommercial space industry of today. Such a vehicle would allow delivery of a scientific payload to a circularized low Earth orbit (LEO) at an altitude of roughly 300 km without requiring the use of expendable stages. These reusable launch vehicles (RLVs) would reduce operational cost in the long term and allow for rapid successive launches. In the 1980s, two British engineers ( and Bob Parkinson) at Rolls Royce began working on the development of a propulsion system that could be utilized in a SSTO vehicle. The result was the RB545 that was envisioned to power the horizontal take-off and landing (HOTOL) vehicle. However, the project development never gained ground due to a lack of availability of other nations to assist in the development of such an extensive program, and the HOTOL project was canceled [1]. The company Ltd. (REL) was formed by Bond and a team of engineers to further develop the technology and produce a design for a viable SSTO vehicle. The engine that has been in development is known as the SABRE, Figure 1, and is intended to power the envisioned aircraft-like SSTO vehicle known as . The SKYLON spacecraft, at the current configuration C1, is designed to takeoff horizontally from a runway using two SABREs operated in air-breathing mode and continue on an ascent trajectory to a pre-determined altitude. The SABREs are then switched over to a more standard rocket mode of operation, followed by adjusting to a more aggressive ascent profile and continuing until the vehicle reaches an apogee of 300 km. The orbit can then be circularized and any scientific payloads aboard can be released into LEO [1]. As seen in Figure A1 of Appendix A, conventional rocket propulsion techniques currently dominate the commercial launch industry due to the high thrust to weight ratio of the propulsion system. However, as observed in Figure A2, the SABRE offers the highest thrust to weight ratio of the air-breathing propulsion methods and is the best candidate for a SSTO vehicle that incorporates air-breathing propulsion. From Figure A2, it is seen that the SABRE has a significantly higher , compared to conventional rocket engines, which results in lower resource expenditure since the air captured during air-breathing mode acts as the oxidizer in the combustion process.

1 Aerospace Engineering Sciences, University of Colorado, Boulder, CO 80309. 1 American Institute of Aeronautics and Astronautics

Figure 1: Synergetic Air-breathing (SABRE). Source: Fig. 1 of Mehta et al. [6]

II. SABRE Operational Overview The operation of the SABRE engine is split into two distinct modes: air-breathing mode and rocket mode. In air-breathing mode, the air intake at the foremost section of the engine ingests freestream air which is then used as the oxidizer that reacts with hydrogen in the combustion chambers. The combustion by- products are then expanded out of the rocket nozzles to generate thrust. At a predetermined altitude and velocity, the conical center body in the air intake translates forward and closes off the system from freestream air [2]. While this sealing process is occurring, the (LOX) pump demand to the is gradually increased to prevent any discontinuity in thrust. The SABRE is then considered to be in rocket mode and utilizes gaseous oxygen and hydrogen for combustion (oxygen must be in the gaseous phase in order to utilize the same injector as is used in air-breathing mode).

III. SABRE Key Components Combining the air-breathing and rocket modes of operation into one system increases complexity of the vehicle. However, the overall efficiency increases due to the absence of a separate inactive system that would otherwise increase the aerodynamic drag on the vehicle. The SABRE design is broken down to the subsystem level in order to discuss the aspects that are vital to the successful performance of the engine. A flow diagram of the system is shown in Figure 2 and is referenced frequently in the subsequent sections.

Air Intake The air intake is comprised of a 20 degree axisymmetric conical center body, with the ability to translate axially, that actively controls the conditions at which freestream air enters the nacelle. During takeoff/subsonic conditions, the center body is fully retracted in order to maximize air intake at lower speeds. As the freestream velocity increases to supersonic conditions, the center body translates forward so that the oblique shock wave that develops at the front of the center body is not ingested by the air intake. The center body is continuously adjusted in order to provide inlet conditions that result in an appropriate pressure recovery value so that the turbo-, shown in Figure 2, is supplied with a nearly constant pressure in order to maximize combustion efficiency. As an aside, the intake is pitched 7 degrees nose-

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down, seen in Figure 1, in order to compensate for the angle of the ascent trajectory of a SSTO vehicle such as SKYLON and maximize inlet air flow [2].

Figure 2: Simplified SABRE Flow Diagram Source: Fig. 5 of Longstaff et al. [1]

Once the optimum conditions for transition from air-breathing to rocket mode have been achieved, predicted to be approximately Mach 5.1 at an altitude of 26 km, the system is switched over to the conventional rocket mode of operation. At this point, the center body translates forward and picks up three conical frustums that seal against the nacelle and isolate the system from freestream air.

Pre-Cooler During air-breathing operation, the stagnation temperature of the air from the intake is in excess of 1223 K and must be cooled prior to contact with the compressor. This elevated temperature is due to the deceleration of the air from supersonic to subsonic velocities through a series of shock waves (one oblique and one normal). The pre-cooler is one of the most innovative components in the SABRE system and is a vital component in successful operation. As seen in Figure 1, in the current design, the pre-cooler system consists of four heat exchangers that operate in parallel to cool the high-temperature inlet air from the intake. Each heat exchanger is comprised of thousands of small diameter cooling tubes, as seen in Figure 3, that are approximately 1 mm in diameter with a wall thickness between 20 to 40 microns [3]. A portion of the air ingested by the intake is redirected by guide vanes (not shown) that force the air to travel radially inward through the heat exchanger matrix. Once through the heat exchanger, another set of guide vanes (not shown) redirects the flow back to an axial flow direction toward the turbo-compressor [9]. The cooling tubes that make up the heat exchanger matrix can vary in material in order to optimize the weight of the system. Cooling tubes near the exterior of the heat exchanger are exposed to higher temperatures and are made from a nickel alloy such as Inconel 718, while tubes at the cooler interior can be made from a lighter aluminum alloy [7]. Additionally, due to the high aerodynamic drag induced by the matrix, the cooling tubes in the heat exchanger are subject to high inward-radial loading and must be sufficiently supported. A lightweight interior structure is used to support the tubes axially and radially and shim-like structures are incorporated throughout the matrix to uniformly transfer the radial load while still allowing for relative motion due to thermal expansion and contraction of the tubes at various temperatures [9]. 3 American Institute of Aeronautics and Astronautics

Figure 3: Simplified Pre-Cooler Heat Exchanger Diagram Source: Reaction Engines Ltd. [11]

Seen in Figure 2, the closed helium loop supplies the heat exchangers with a heat-sync as the helium, which can be as low as 30 K, circulates through the cooling tubes. The helium is cooled prior to entering the pre-cooler by passing through a heat exchanger with (LH). As an aside, helium is chosen as the coolant, as opposed to hydrogen, in order to avoid hydrogen embrittlement, remain in gaseous phase at temperatures as low as 30 K, and utilize the high ratio of specific heats to reduce the pressure ratio in the closed Brayton cycle loop [2]. Mock heat exchangers have been shown to be capable of cooling the stagnation temperature of the air from 1223 K at the inlet to below 133 K at the interior in 0.01 s [10]. During this rapid cooling process, the closely spaced heat exchanger matrix is vulnerable to clogging from the condensation and subsequent freezing of water that originated as moisture in the inlet air. A unique and innovative solution for the prevention of ice formation in the heat exchanger was made public in 2015 when REL published a patent on the technology. The system utilizes a sophisticated injection system of methanol as an antifreeze agent and a moisture capturing system in order to mitigate the issue [9]. The ice-prevention aspect of the heat exchanger is discussed in detail in Appendix B.

Turbo-Compressor and Helium Turbine Once the inflow air exits the pre-cooler section, it is directed towards the turbo-compressor. The low temperature air is typically at 130 kPa when exiting the pre-cooler and requires the high pressure ratio of the turbo-compressor (around 150:1) in order for the inlet air to be boosted to the range of pressures typically observed in rocket combustion chambers [1]. The turbo-compressor is driven by a specially designed counter-rotating turbine in order to compensate for the difference in the of the hot helium in the turbine versus the cold air in the compressor [11]. The high temperature helium in the Brayton cycle loop is supplied by the addition of heat from the pre- burner section. Fuel rich combustion occurs in the pre-burner section in order to add heat to the helium loop so that the helium turbine is supplied with helium at a relatively fixed temperature. The fuel rich combustion by-products then move to the combustion chamber where the remaining hydrogen reacts with the air in the combustion chamber.

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Bypass As mentioned previously, LH is used to extract heat from the helium before it enters the pre-cooler system. The LH is then diverted to the bypass region in the aft section of the engine to be utilized in the array of ramjets. High temperature air in the bypass stream, that was not diverted through the heat exchangers in the pre-cooler segment, is used as the oxidizer that reacts with the hydrogen injected into the ramjets. Complexity and weight is added to the overall system, however, the drag penalty incurred from simply spilling the bypass air through the rear of the engine is more severe and thus necessitates the use of the array [2].

Nozzle and Combustion Chamber In the current design, a conventional bell contour nozzle is used to accelerate combustion gases to generate thrust. In air-breathing mode, excess hydrogen from the heat exchanger in the helium loop can also be used, in parallel to the ramjet operation, for film cooling of the combustion chamber and nozzle. During rocket mode, both the hydrogen and oxygen must be in gaseous phase, prior to injection into the combustion chamber, in order to utilize the same injector as the air-breathing mode. Similar to the air- breathing mode, the injected hydrogen is maintained in the gaseous phase through fuel rich combustion in the pre-burner, now set to a much lower demand however. The LOX must also be converted to the gaseous phase prior to injection to the combustion chamber. This phase change is achieved through circulation of the oxidizer within the cooling jacket surrounding the main chamber prior to injection to the combustion chamber [2]. As with all fixed geometry nozzles, the thrust generated by the system is negatively impacted when operated at an altitude other than that which the nozzle has been optimized for. At lower altitudes, the nozzle flow is over-expanded, while at higher altitudes, the nozzle flow is under-expanded. To combat this issue, REL has experimented with the development of an expansion deflection pressure compensation nozzle [3]. The expansion deflection nozzle design features a center body that translates axially, based on the atmospheric pressure, and varies the throat area of the nozzle, thus changing the chamber pressure and expansion of the flow in the nozzle. Incorporation of such a system would allow for near-perfect expansion of the flow over a much wider range of operation and increase engine performance.

IV. System Performance and Implementation As previously mentioned, the SABRE system is currently being developed with the purpose of powering the SSTO vehicle, SKYLON. An artist impression of the SKYLON vehicle is shown in Figure 4. The theoretical values shown in Table 1 provide a baseline performance estimate that can be utilized in order to determine whether or not the propulsion system would be capable of powering a SSTO vehicle such as the SKYLON. The equivalence ratio greater than unity indicates that more fuel is used by the system than what

Figure 4: Artist Impression of SKYLON Vehicle Source: Fig. 2 of Longstaff et al. [1] 5 American Institute of Aeronautics and Astronautics

is required for stoichiometric combustion. This excess fuel is due to the use of LH to extract heat from the helium loop and the subsequent utilization of the hydrogen in the bypass array and nozzle film cooling applications. From Table 1, the difference in specific impulse (퐼푠푝) is evident between the two operational modes. However, the value calculated for the air-breathing mode far exceeds the predicted maximum depicted in Figure A2. This may be due to an unspecified mass flow rate of methanol in the ice-prevention system that is expelled through the combustion chamber and/or an undocumented amount of LOX that may be utilized in the pre-burner system to add heat to the helium Brayton cycle loop. The 퐼푠푝 in rocket mode appears slightly higher than the average for LH/LOX liquid engines. This may further support the previous observation that pressure accommodation nozzles are being explored by REL, which could in-turn increase the thrust coefficient (퐶퐹) of each nozzle.

Table 1: Performance Parameters of the SABRE2 [1]

Air-Breathing Mode Rocket Mode Parameter (Max Thrust) (100% Throttle) Hydrogen Mass Flow 31 kg/s 46 kg/s Air Mass Flow 382 kg/s - LOX Mass Flow - 278 kg/s Chamber Pressure (nominal) 10.38 MPa 14.5 MPa Thrust (core only) 1,045 kN - By-Pass Thrust 405 kN - Total Thrust 1,450 kN 1,458 kN Specific Impulse 4,772 s 458 s Equivalence Ratio3 2.8 - Altitude 20,270 m - Mach No. 4 - Intake Pressure Recovery Factor 0.1592 -

Additional details of the SABRE and SKYLON vehicle are shown in Table 2 and Table 3. The inert mass fraction (푓푖푛푒푟푡) of the system is large, 0.21, compared to the typical value for conventional rockets near 0.1. The added mass is due to the increased complexity of the SABRE system which must be carried through the full ascent profile. Significant design effort is required in order to decrease 푓푖푛푒푟푡 to an acceptable range. Component design margins and lifetimes must be reduced in parallel to the implementation of innovative and lightweight structural solutions in order to achieve a significantly lower inert mass.

Table 2: Projected SABRE Design Characteristics

Parameter Predicted Value Overall Length 13 – 15 m Nacelle Diameter 3 – 4 m Total Mass (including nacelle 7,482 kg structure) Thrust to Weight Ratio4 ~ 20

2 Performance is defined for a single Nacelle, defined as the overall system depicted in Figure 1. 3 Ratio of fuel utilized in core system divided by the amount used for stoichiometric combustion. 4 Calculated value does not match data shown in Figure A2; total mass may have been under-reported. Additional mismatch in location of maximum thrust with respect to noted as well. 6 American Institute of Aeronautics and Astronautics

Table 2: Projected SKYLON C1 Design [1]

Parameter Predicted Value Overall Length 84 m Wing Span 25 m Fuel Mass 66,699 kg Oxidizer Mass 149,931 kg Total Propellant Mass 216,630 kg Inert Mass5 57,170 kg Payload Mass 10,275 kg Takeoff Mass 275,000 kg Inert Mass Fraction 0.21

The current launch profile for the SKYLON C1 vehicle begins with an ascent to 26 km in altitude over a period of 694 seconds, at which point the vehicle is traveling at Mach 5.1 and is 620 km downrange. A change in the propulsion method is initiated and switches the SABREs to rocket mode. The trajectory steepens and the vehicle reaches main engine cut-off (MECO) after 285 seconds at an altitude of 80 km. Residual propellants are then jettisoned and the vehicle coasts for 44 minutes before reaching an apogee of 300 km, at which point the orbit can be circularized using onboard thrusters. Payload doors are opened and the payload is inserted into LEO. Retrograde thrusters are then fired onboard the SKYLON to decrease orbital velocity and attain a re-entry profile leading back to a suitable landing location. The vehicle is unpowered during the descent segment and vehicle momentum is managed via a series of S turns [1]. Several plots of the trajectory are shown in Figure A3.

V. Feasibility and Future Design Changes Proving the capability of the pre-cooler technology that is utilized in the SABRE design is a vital step in the development program. Small scale testing of the heat exchanger technology has demonstrated the thermal capabilities as well as the effectiveness of the ice-prevention system. Full scale testing of the propulsion system, including the pre-cooler, is expected to commence in 2020 and will further validate the technology. The implementation of the SABRE system as the propulsion method for a SSTO vehicle, however, brings about challenges that require attention. Due to the relatively high mass of the SABRE compared to its conventional counterparts, weight savings must be achieved in other areas of the vehicle. For the SKYLON, this has driven the need for research into lightweight design solutions for the structure of the vehicle to reduce the inert mass. With the increase in the percentage of unproven technologies within the system, the development phase is significantly extended. The current C1 configuration of the SKYLON vehicle incorporates a payload mass of 10,275 kg, with a target mass being 12,000 kg. However, in 2010 the SKYLON requirement specifications were revised to reflect the current commercial launch market. This resulted in an increase of the payload mass to 15,000 kg and enlarged the payload bay to 13 m long and 4.8 m in diameter. The changes resulted in a total increase of the gross takeoff weight (GTOW) by about 20% for the final D1 configuration of the SKYLON vehicle. Aside from the vehicle weight issue, the implementation of the SABRE system into a SSTO vehicle has the potential to disrupt the of the vehicle. The SABREs are required to be placed sufficiently far forward on the vehicle to ensure stability over the entire flight regime. In this position, the center of gravity (CG) is maintained upstream of the center of pressure (CP) during powered flight while the center of gravity shifts due to propellant consumption. However, with the SABREs positioned at

5 After takeoff, 1200 kg of water for the emergency brake cooling system is vented overboard since no longer required. 7 American Institute of Aeronautics and Astronautics

approximately the midpoint of the vehicle, there exists the potential for the plumes from the engines to disrupt the boundary layer flow along the fuselage. Figure 5, from Mehta et al. (2015), shows contours of the stagnation temperatures throughout the domain of a SKYLON D1a vehicle in powered flight at various speeds. It should be noted that the simulation is generated using the inviscid Euler approach and does not accurately depict shock interactions or boundary layer flows. However, this modeling approach is sufficient for the purpose of identifying the presence of the issue. As seen, the plume from the SABRE grows as the vehicle gains altitude and velocity until contact with the fuselage. This interaction with the fuselage is highly likely to cause plume induced flow separation (PIFS) along the fuselage and generate excessive heating in the tail region of the vehicle [6]. Further analysis with more appropriate computational methods is required to accurately quantify the temperature increases in the regions of interest. However, it is evident that the SABREs in the current design configuration threaten the structural integrity in the tail region of the vehicle. One viable solution to this issue may be found in the experiments that REL is conducting with pressure compensating nozzles. If an expansion deflection nozzle is developed and can perfectly expand the flow at higher altitudes, the width of the plume from the SABRE would be significantly reduced and not pose a threat to the tail section of the vehicle.

Figure 5: Stagnation Temperature Contours of SKYLON Vehicle During Powered Flight (a) 푀∞ = 6.673, ℎ = 39.695 푘푚, (b) 푀∞ = 8.577, ℎ = 51.983 푘푚 (c) 푀∞ = 12.189, ℎ = 65.160 푘푚, (d) 푀∞ = 16.969, ℎ = 75.771 푘푚 Source: Fig. 15 of Mehta et al. [6]

VI. Conclusion The SABRE is a hybrid propulsion system that utilizes both the concept of a conventional engine as well as a rocket engine throughout the operating regime. The system is a potential candidate for powering a SSTO vehicle such as the SKYLON, but imposes design constraints on the system it is integrated into. Due to the mass of the SABRE system, the SSTO vehicle must incorporate lightweight design solutions to compensate for the weight of the propulsion system. The exact operational

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performance of the SABRE is unclear at this point, however, sensitivity to atmospheric conditions during air-breathing mode of powered flight has the potential to create narrow performance margins that negatively impact the viability. With the current mass of the SABRE, SSTO vehicles would be required to mount the system further forward which increases the potential for the plume to affect the vehicle structure. To mitigate the issue, the SABRE system may be required to incorporate pressure compensating nozzles to reduce the plume size at higher altitudes. Looking at the SKYLON as an example, the increase of GTOW in the newest D1 configuration stretches the capability of the SABREs. This fact necessitates an increase in power density of the SABRE system in order to make the SABRE a more viable candidate for powering a SSTO vehicle such as the SKYLON.

References [1] Longstaff, R., Bond, A., “The SKYLON Project,” AIAA Journal, DOI: 10.2514/6.2011-224. [2] Varvill, R., Bond, A., “The SKYLON ,” JBIS, Vol. 57, pp. 22-32, 2004. [3] Varvill, R., Bond, A., “The SKYLON Spaceplane: Progress to Realisation,” JBIS, Vol. 61, pp. 412-418, 2008. [4] Varvill, R., Bond, A., “A Comparison of Propulsion Concepts for SSTO Reusable Launchers,” JBIS, Vol. 56, pp. 108-117, 2003. [5] Burns, B., “HOTOL space transport for the twenty-first century,” Proc. Instn. Mech. Engrs., Vol. 204, 1990. [6] Mehta, U., Aftosmis, M., Bowles, J., Pandya, S., “Skylon Aerodynamics and SABRE Plumes,” AIAA Journal, DOI: 10.2514/6.2015-3605. [7] Varvill, R., “Heat Exchanger Development at Reaction Engines Ltd.,” IAC-08-C4.5.2. [8] Bond, A., Varvill, R., “Engine,” UK Patent GB 2519155 B, issued date Oct. 12, 2016. [9] Bond, A., Varvill, R., “Heat Exchangers,” UK Patent GB 2519147 A, issued date April 15, 2015. [10] “SABRE (rocket engine),” Wikipedia Available: https://en.wikipedia.org/wiki/SABRE_(rocket_engine). [11] “SABRE,” Reaction Engines Available: https://www.reactionengines.co.uk/sabre/.

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Appendix A – Supplemental Images

Figure A1: Potential Propulsion Methods for SSTO Vehicles Source: Fig. 1 of Burns [5]

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Figure A2: Thrust/Weight Ratio and Specific Impulse of SSTO Propulsion Systems Source: Fig. 7 and 8 of Varvill et al. [4]

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Figure A3: Predicted Ascent and Descent Profiles of SKYLON C1 Vehicle Source: Fig. 8 of Longstaff et al. [1]

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Appendix B – Heat Exchanger Ice-Prevention System

The following information was extracted from the patent that REL filed in 2015, [9], on the ice- prevention technology that the heat exchanger in the SABRE uses. Figure B1 shows the heat exchanger with the spiral arrangement of cooling tubes. The view is shown in the axial direction and the circles depicted at various points in the figure can be interpreted as rods or tubes extending in and out of the plane. Methanol is injected at the outermost section and mixes with the condensed water droplets that are entrained in the flow. The methanol that is mixed with the water lowers the freezing temperature and prevents the formation of ice as the liquids condense and are collected in the moisture capturing system. Air flows radially inward and caries the liquid droplets towards the moisture capturing systems, referred to as ‘catchers’, as seen in Figure B2. The first row of catcher rods disrupts the incoming flow and funnels the flow towards catcher rods located in the second row. Each catcher rod is comprised of a silica coated mesh surface that encapsulates each rod and extracts the liquid droplets, on contact, from the airstream via surface tension. The liquid droplets then move to collection cavities where the liquid can be extracted through small scavenge holes via a suction pump. From testing, it is observed that the first row only contributes 5% of the total amount of liquid that is captured by the system. This means that in future iterations, the first row could be replaced with solid rods to act as bluff bodies that would perform the same function in terms of funneling flow onto the catcher rods in the second row for moisture collection. Figure B1 shows only one moisture capturing sequence (outer injector array and two levels of catcher assemblies). However, if several of these systems are stacked radially outward, there would be several regions of moisture capturing throughout the heat exchanger matrix. In this scenario, to utilize the methanol efficiently, the methanol-water mixture is recycled in a unique counter-flow manner. At the inner most moisture capturing segment, pure methanol is injected and flows inward and is captured, along with a small amount of water from the air. This methanol-water mixture, primarily methanol, is routed to an injection site further upstream (radially outward) from the first injection site. The mixture then flows inward and is collected by the catchers just downstream (radially inward). The liquid that is collected at this site now has a slightly higher concentration of water. The higher concentration of water is acceptable since the collection site is at a higher temperature, being that it is located further upstream in the heat exchanger matrix. This process of diverting the collected mixture further and further outward is repeated until the last segment where the collected liquid can be expelled through the combustion chamber. This counter-flow process of the methanol allows for efficient use and minimizes consumption of the anti-freeze agent.

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Figure B1: Axial View of Heat Exchanger Methanol injection site (184), cooling tube (120), moisture capturing system (detail E) Source: Fig. 5A and 19C of [9]

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Figure B2: Detailed View of Moisture Capturing System Incident liquid droplets in air (520), silica coated mesh (258), collection cavities (262), liquid scavenge holes (266). Source: Fig. 19D and 13I of [9]

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