On a New Type of Combined Solar–Thermal/Cold Gas Propulsion System Used for LEO Satellite’S Attitude Control

applied sciences

Article On a New Type of Combined Solar–Thermal/Cold Gas Propulsion System Used for LEO Satellite’s Attitude Control

1 1 2, 1 1 Constantin Sandu , Valentin Silivestru , Grigore Cican *, Horat, iu S, erbescu , Traian Tipa , Andrei Totu 1 and Andrei Radu 1 1 National Research and Development Institute for Gas Turbines COMOTI, 220D Iuliu Maniu, 061126 Bucharest, Romania; [email protected] (C.S.); [email protected] (V.S.);

[email protected] (H.S, .); [email protected] (T.T.); [email protected] (A.T.); [email protected] (A.R.) 2 Faculty of Aerospace Engineering, Politehnica University of Bucharest, 1-7 Polizu Street, 011061 Bucharest, Romania * Correspondence: [email protected]

Received: 15 September 2020; Accepted: 12 October 2020; Published: 15 October 2020

Abstract: This paper presents the development, construction and testing of a new type of solar–thermal propulsion system which can be used for low earth orbit (LEO) satellites. Currently, the vast majority of LEO satellites are fitted with a cold gas propulsion system. Although such a propulsion system is preferred, the service duration of an LEO satellite is limited by the amount of cold gas they carry onboard. In the case of the new type of solar–thermal propulsion system proposed in this paper, the cold gas is first transferred from the main tank in a cylindrical service tank/buffer tank which is placed in the focal line of a concave mirror. After the gas is heated by the solar light focused on the service tank by the concave mirror, it expands by opening the appropriate solenoid valve for the satellite’s attitude control. In this way the service duration of LEO satellite on orbit can increase by 2.5 times compared with a classic cold gas propulsion system. This is due to the propellant’s internal energy increase by the focused solar light. This paper also presents the results achieved by carrying out tests for the hot gas propulsion system in a controlled environment.

Keywords: solar energy; solar–thermal; cold gas; propulsion system; LEO satellites

1. Introduction Since the first satellite, Sputnik, was launched in 1957, more than 5560 successful launched operations occurred, which means about 9600 satellites in their orbits around the Earth’s. The European Space Agency (ESA)’s Space Debris Office estimated that about 5500 of these satellites were in space until February 2020, while 2300 were still functioning [1], most of which are low earth orbit (LEO) satellites. These satellites operate fairly close to the surface of the Earth, between 500 and 2000 km. Because the number of artificial satellites in the Earth’s orbit is increasing, their life-time must increase so that the quantity of space debris will not increase in a way that will affect space security, i.e., collision between satellites or spacecraft [2]. A satellite’s operational life is measured also taking into account the quantity of propellant taken onboard and utilized for the orbit-position correction system, attitude control and specific maneuvers on the orbit. The efficiency of these correction systems depends on the propulsion system type. The most common types are cold gas propulsion systems, chemical propulsion, electric propulsion, propellantless propulsion systems—solar sails and solar thermal propulsion, etc.

Appl. Sci. 2020, 10, 7197; doi:10.3390/app10207197 www.mdpi.com/journal/applsci Appl. Sci. 2020, 10, 7197 2 of 18

More details about this propulsion systems can be found in [3–5]. There is an increasing effort to develop correction systems with an increased efficiency by utilizing energy outside the satellite, e.g., the solar energy for solar sails [6,7] or solar energy for solar–thermal propulsion [8]. Furthermore, the operational life increase of the satellite leads to environmental protection, the saving of energy, materials, manpower and finally, cost saving. The use of solar energy in the cosmic space was imagined for propulsion systems fitted with a mirror [9] but also for other activities, e.g., space solar power satellite [10] or a specific surface illumination [11]. The idea of a mirror in space to reflect sunlight and thus generate power and light on Earth was proposed in 1928 by Hermann Oberth who postulated a space-manufactured 5 mm-thick mirror using sodium for the reflective layer, orbiting Earth at a 1000–5000 km altitude [12]. Across time, various solar concentration schemes have been proposed, including lenses, deployable arrays and inflatable arrays [13–15]. A very comprehensive conceptual, technical and socio-economic study and the exhibition of space mirrors was conducted by the space visionary Krafft Ehricke [16]. He proposed and analyzed in some detail a number of generic applications for providing lunar-type night illumination service (“Lunetta”), solar-type light energy services (“Soletta”), insolation for bio-production enhancement (“Biosoletta”) to produce food and biomass, insolation for agricultural weather stabilization, precipitation management, crop drying and desalination (“Agrisoletta”) and insolation for generating electricity on earth (“Powersoletta”). A significant brief step in the development of space mirrors was the Russian Space Mirror Project “Znamya” (banner) developed by the “Space Regatta Consortium” (SRC) established in 1990 by the Russian Space Agency and the corporation Energia [17]. Detailed information about the Znamya experiments is somewhat sparse [18]. Research on solar–thermal propulsion has been conducted for more than 60 years [19], yet no flight testing has been achieved. An overview of the milestones and program associated with solar–thermal technology since the advent of the technology in 1956 is provided [20]. The solar–thermal propulsion concept is a fairly basic one, relying on highly concentrated sunlight to heat a monopropellant to very high temperatures (typically approaching 3000 K) and subsequently exhausting the propellant through a discharge nozzle to provide thrust. A typical solar–thermal engine is composed of three primary components or subassemblies:

A solar concentrator, which concentrates and focuses solar energy onto the receiver; • The receiver itself, which acts as a heat exchanger for the propellant; • The propellant storage and feed system, which includes tankage and feed lines for routing the • propellant to the receiver. A fourth component, control electronics, is not considered here but it is required to retrieve • system telemetry, open and close valves, and (potentially) control both the concentrator and receiver position.

A working principle scheme of a solar–thermal propulsion system [21], further called STP, is presented in Figure1. A beneﬁt of using STP as a spacecraft propulsion system is that it uses solar radiation directly to heat the propellant, therefore improving the conversion eﬃciency over resisto-jets that require converting solar radiation into electrical energy and then into thermal energy to heat the propellant. The operating temperature is limited by the materials available. A cold gas propulsion system (CGP) relies on the process of controlled ejection of compressed liquid or gaseous propellants to generate thrust. Due to the absence of a combustion process, a CGP system requires only one propellant (without an oxidizer), and hence can be designed with minimal Appl. Sci. 2020, 10, x FOR PEER REVIEW 3 of 18

Appl. Sci. 2020, 10, 7197 3 of 18 complexity. The schematic of a typical CGP system is shown in Figure2, and the main components include a propellant storage and a nozzle [22]. Appl. Sci. 2020, 10, x FOR PEER REVIEW 3 of 18

Figure 1. Schematic of a solar thermal propulsion system.

A benefit of using STP as a spacecraft propulsion system is that it uses solar radiation directly to heat the propellant, therefore improving the conversion efficiency over resisto-jets that require converting solar radiation into electrical energy and then into thermal energy to heat the propellant. The operating temperature is limited by the materials available. A cold gas propulsion system (CGP) relies on the process of controlled ejection of compressed liquid or gaseous propellants to generate thrust. Due to the absence of a combustion process, a CGP system requires only one propellant (without an oxidizer), and hence can be designed with minimal complexity. The schematic of a typical CGP system is shown in Figure 2, and the main components include a propellant storage and a nozzle [22]. Figure 1. Schematic of a solar thermal propulsion system.

The simplersimpler design design of aof CGP a CGP system system leads toleads a smaller to a systemsmaller mass system and lowermass powerand lower requirements power requirementsfor regulation purposes.for regulation However, purposes. these advantagesHowever, comethese at advantages the cost of a monotonicallycome at the decreasingcost of a monotonicallythrust profile over decreasing a period ofthrust time. profile The thrust over produced a period is of directly time. proportionalThe thrust produced to the pressure is directly of the proportionalpropellant inside to the the pressure tank (propellant of the propellant storage) inside and overthe tank the course(propellant of the storage) mission, and the over tank the pressure course ofdecreases the mission, (due tothe propellant tank pressure usage) decreases resulting (due in a to decrease propellant in the usage) maximum resulting thrust in a generated decrease byin the maximumsystem. There thrust are generated many studies by the and system. applications There aboutare many cold studies gas systems, and applications CGPs [23–26 about]. cold gas systems,The CGPs novelty [23–26]. element of the present paper consists in the fact that it proposes a solution which combinesThe novelty the solar–thermal element of propulsion the present and paper cold consists gas propulsion in the fact concept, that soit proposes that the solar a solution energy which which is concentrated from a concave mirror will directly heat up the propellant tank. The propellant tank is combines the solar–thermalFigure propulsion 2. Schematic and of cold the cold gas gaspropulsion propulsion concept, system. so that the solar energy whichplaced is in concentrated the mirror’s from focal a axis. concave mirror will directly heat up the propellant tank. The propellant tank Theis placed presentsimpler in the paper design mirror’s tries of to focala solveCGP axis. the system operational leads lifeto issuea smaller by developing system mass a propulsion and lower system power for satellitesrequirementsThe present which for will paper regulation increase tries to the solvepurposes. operational the operational However, life of the life these satellite issue byadvantages on developing the orbit. come a propulsion at the systemcost of for a satellitesmonotonicallyTests which at low decreasingwill temperature increase thrust the (specific operational profile for coldover life gas aof period propulsion)the satellite of time. andon the testsThe orbit. atthrust high produced temperature is (specificdirectly forproportional hotTests gas at propulsion) low to the temperature pressure will be of (specific carried the propellant out. for cold A comparison inside gas propulsion) the tank between (propellant and the test forcess at storage) high obtained temperature and in over both the (specific cases course for fordiof ff thehoterent mission, gas pressures propulsion) the willtank will bepressure carried be carried decreases out. out. The A aim(due comp is to toarison propellant see what between improvement usage) the resultingforces in obtained performance in a decrease in both will incases the forhotmaximum different gas propulsion thrust pressures generated system will bring,be by carried the given system. out. the The sameThere ai amountm are is tomany see of masswhatstudies ofimprovement Nand2 used applications in both in performance the about cold cold and will gas hot gas propulsion systems. thesystems, hot gas CGPs propulsion [23–26]. system bring, given the same amount of mass of N2 used in both the cold and hot gasThe propulsion newnovelty type element of systems. solar–thermal of the present propulsion paper system consists developed in the fact by COMOTIthat it proposes in the frame a solution of a national which spacecombines research the solar–thermal program STAR—project propulsion STRAUSS—canand cold gas propulsion be a solution concept, for the so satellite’sthat the solar operational energy lifewhich increase. is concentrated from a concave mirror will directly heat up the propellant tank. The propellant tank is placed in the mirror’s focal axis. The present paper tries to solve the operational life issue by developing a propulsion system for satellites which will increase the operational life of the satellite on the orbit. Tests at low temperature (specific for cold gas propulsion) and tests at high temperature (specific for hot gas propulsion) will be carried out. A comparison between the forces obtained in both cases for different pressures will be carried out. The aim is to see what improvement in performance will the hot gas propulsion system bring, given the same amount of mass of N2 used in both the cold and hot gas propulsion systems. Appl. Sci. 2020, 10, x FOR PEER REVIEW 4 of 18

The new type of solar–thermal propulsion system developed by COMOTI in the frame of a national space research program STAR—project STRAUSS—can be a solution for the satellite’s Appl. Sci. 2020 10 operational, life, 7197increase. 4 of 18

2. New Type of Solar–ThermalSolar–Thermal/Co/Coldld Gas Propulsion DescriptionDescription The operating principle of this new propulsion system system is is based based on on the the focus focus of of solar solar light. light. According to this concept, the working life of thrustersthrusters can be increased if additional energy is taken from Sun for the propellant’s internalinternal energyenergy increaseincrease beforebefore itsits expansionexpansion inin thethe nozzles.nozzles. The new propulsion system is composed of a concaveconcave mirror having a small service propellant tank placed along its focal line (Figure3 3).).

Figure 3. Hot gas propulsion system working principle.

The coldcold gas gas (nitrogen, (nitrogen, N 2)N is2) transferredis transferred from from the main the tankmain with tank a speciﬁcwith a timespecific before time expansion. before Theexpansion. service tankThe service is a small tank cylinder is a small made cylinder of stainless made steel of whichstainless is paintedsteel which in black is painted for a total in absorptionblack for a oftotal solar absorption light. When of solar the solar light. light When is focused the solar on theligh servicet is focused propellant on the tank, service its temperature propellant increasestank, its withtemperature several hundredincreases degreeswith several Celsius, hundred thus heating degrees the Celsius, gas inside. thus heating the gas inside. During expansion, the propulsion force is higher than in the case of a cold gas because the expansion speed is higher. In this way, an important quantity of gasgas is saved at every gas expansion for attitudeattitude correction.correction. TheThe obtained obtained reaction reaction force force is suis ﬃsufficientcient to changeto change the the satellite’s satellite’s attitude, attitude, so that so itthat maintains it maintains its correct its correct position position in space. in Byspace. heating By upheating the propellant up the propellant inside the inside service the propellant service tank,propellant more correctionstank, more of corrections attitude/gas of expansions attitude/gas than expansions in the case than of classic in the cold case gas of propulsion classic cold system gas canpropulsion be achieved. system can be achieved.

3. Experimental System Description In order to proveprove thethe workingworking principleprinciple of thethe system,system, experimentalexperimental equipment was built for laboratory testing.testing. ForFor testing,testing, thethe radiation radiation source—which source—which in in real real life life is is the the sun—was sun—was replaced replaced with with an infraredan infrared lamp lamp system system which which emits emits an equivalent an equivale radiationnt radiation with thewith one the present one present on the on LEO. the The LEO. design The ofdesign the experimentalof the experimental equipment equipment destined destined for the for testing the testing of the of propulsionthe propulsion principle principle is presented is presented in Figurein Figure4. 4. The testing equipment consisted of a light source (1) which was an infrared lamp system—for generating the parallel infrared rays. The infrared rays were focused by the mirror (2) on the central focal line. The service propellant tank (3) was mounted on the central focal line of the mirror, therefore it was heated up. The cold N2 gas was fed from the main tank (4) which had N2 at 200 bar into the pressure regulator (5). The pressure regulator reduced the gas pressure to 20 bar and transferred the gas with this pressure to the service propellant tank. When the solenoid (6) valve opened brieﬂy (t = 0.1–0.4 s), a certain a quantity of gas was exhausted through the Laval nozzle (7) generating thrust force. The thrust force was measured by the force transducer (8). Five thermocouples were used for the temperature measurement; one mounted on the propellant tank (9a), three used for the heat dissipation measurement along the distance from Appl. Sci. 2020, 10, 7197 5 of 18 the radiation source to the mirror (9b), and one thermocouple was used for the room temperature measurementAppl. Sci. 2020, 10, (9c). x FOR PEER REVIEW 5 of 18

Figure 4. The design ( left) and photo ( right) of the experimental equipment.

The pressuretesting equipment transducer consisted was mounted of a onlight the source service (1) propellant which was tank. an For infrared measuring lamp the system—for irradiation, onegenerating pyrometer the parallel (10) was infrared used. rays. The infrared rays were focused by the mirror (2) on the central focal Theline. values The service of the propellant gas pressure, tank temperature (3) was mounted and thrust on the force central were focal transmitted line of the to amirror, data acquisition therefore systemit was heated (11). up. The cold N2 gas was fed from the main tank (4) which had N2 at 200 bar into the pressureFurthermore, regulator the(5). solarThe pressu constantre regulator at the mean reduced distance the gas of the pressure Earth fromto 20 bar the Sunand transferred was measured the togas be with S = this1366 pressure W/m2 (with to the an service uncertainty propellant of 3 tank. W/m 2)[27]. Therefore, the radiation power used in 2 ± 2 the experimentsWhen the solenoid was PR =(6)1360 valve W /mopened. The briefly infrared (t light= 0.1–0.4 source s), isa capablecertain ofa quantity emitting 5000of gas W was/m , beingexhausted powered through with the 220 Laval V at nozzle 60 Hz. (7) A generating potentiometer thrust was force. used The to thrust control force the was intensity measured of the by light. the Theforce radiation transducer source (8). Five was thermocouples placed at 1 m above were theused focal for axisthe temperature of the mirror. measurement; one mounted on theThe propellant irradiance tank value (9a), was three recorded used byfor the the pyranometer. heat dissipation Convection measurement is negligible along inthe space, distance given from the absencethe radiation of su ﬃsourcecient gasto the molecules mirror (9b), for energy and one transportation. thermocouple Therefore, was used onlyfor the an room approximation temperature for energymeasurement loss through (9c). radiation (due to mirror heating) can be done. The radiation losses are determined withThe the Stefan–Boltzmannpressure transducer law was [28], mounted see Equations on the (2) andservice (3) belowpropellant in the tank. article. For measuring the irradiation,The solenoid one pyrometer valve functions (10) was betweenused. 10 C and 125 C with air or nitrogen. The shortest − ◦ ◦ openThe/close values time intervalof the isgas 0.02–0.03 pressure, s. Ittemperature can operate betweenand thrust 0 and force 90 barwere and transmitted is powered to with a 24data V A.C.acquisition current system [29]. (11). Furthermore, the solar constant at the mean distance of the Earth from the Sun was measured to 3.1.be S The= 1366 Mirror W/m2 (with an uncertainty of ±3 W/m2) [27]. Therefore, the radiation power used in the experimentsThe main was important PR = 1360 piece W/m of2. theThe equipmentinfrared light isthe source concave is capable mirror. of Theemitting mirror 5000 calculation W/m2, being was carriedpowered out with as in220 [30 V]. at Design 60 Hz. and A dimensionspotentiometer of thewas mirror used canto control be seen the in Figureintensity5. of the light. The radiationThe openingsource was of theplaced mirror at 1/ aperturem above istheL focal= 904 axis mm, of mirrorthe mirror. width is l = 410 mm, and the focal line heightThe irradiance is h = 268 value mm was from recorded the mirror by base.the pyranometer. The parabola Convection parameter is is negligiblef = p/2 = 268 in space, mm and given its Equationthe absence is givenof sufficient by (1): gas molecules for energy transportation. Therefore, only an approximation for energy loss through radiation (due to mirror heating) can be done. The radiation losses are determined with the Stefan–Boltzmannx2 lawx2 [28], see 452Equations2 (2) and (3) below in the article. y = = = = 190.58 mm (1) The solenoid valve functions4 f between4 268 −10 °C4 and268 125 °C with air or nitrogen. The shortest × × open/close time interval is 0.02–0.03 s. It can operate between 0 and 90 bar and is powered with 24 V A.C. current [29].

3.1. The Mirror The main important piece of the equipment is the concave mirror. The mirror calculation was carried out as in [30]. Design and dimensions of the mirror can be seen in Figure 5. Appl. Sci. 2020, 10, x FOR PEER REVIEW 6 of 18

Appl. Sci. 2020, 10, 7197 6 of 18 where p is the parabola parameter which can be expressed as p = 2f. The parabola Equation can be expressed as x2 = 2py; y is the vertical Cartesian coordinate in which the parabola is represented, x is the horizontal Cartesian coordinate, f is the ordinate of focus point, F. For the parabola: F(0, f ) when the parabola Equation is given by Equation: y = (1/4f )x2. Replacing x with half of the aperture and f with the height of the focal line from the mirror base, we obtain the height of the mirror, y as in Figure6, withAppl. theSci. 2020 value, 10,of x FOR 190.5 PEER mm. REVIEW 6 of 18 Figure 5. The design and photo of the mirror of the experimental equipment.

The opening of the mirror/aperture is L = 904 mm, mirror width is l = 410 mm, and the focal line height is h = 268 mm from the mirror base. The parabola parameter is f = p/2 = 268 mm and its Equation is given by (1): 𝑥 𝑥 452 𝑦 = = = = 190.58 mm (1) 4𝑓 4 × 268 4 × 268 where p is the parabola parameter which can be expressed as p = 2f. The parabola Equation can be expressed as x2 = 2py; y is the vertical Cartesian coordinate in which the parabola is represented, x is the horizontal Cartesian coordinate, f is the ordinate of focus point, F. For the parabola: F(0,f) when the parabola Equation is given by Equation: y = (1/4f)x2. Replacing x with half of the aperture and f with the height of the focal line from the mirror base, we obtain the height of the mirror, y as in Figure 6, with the valueFigureFigure of 190.5 5.5. The mm. design and photo of the mirror mirror of of the the experimental experimental equipment. equipment.

Figure 6. Parabola Equation.

The focus point is the converging point of all thethe reflectedreflected rays when the incident rays are in parallel with the OY axis of the coordinates. The mi mirrorrror is made of carbon fiber-reinforced fiber-reinforced polymer, CFRP. ItIt hashas aa totaltotal thicknessthickness ofof 77 mmmm andand hashas ribsribs forfor stistiffnessffness onon thethe oppositeopposite sideside ofof thethe mirror’smirror’s reflectivereflective face.face. In order toto buildbuild thethe concaveconcave mirror,mirror, aa specificspecific moldmold waswas required.required. The mold was made out of cast iron (Figure7 7).). It features a fine polished concave side according to Equation (1) and with the same dimensions as presented in Figure5 for the concave mirror. Multiple layers of pre-impregnated carbon fiber with a thickness of 0.3 mm were used for the concave mirror manufacturing.

During this process, the cast iron mold was covered with resin-absorbing cloth, placed in a Figure 6. Parabola Equation. vacuum bag which was vacuumed at p = 0.01 bar and treated in oven at 180 ◦C for 6 h. The focus point is the converging point of all the reflected rays when the incident rays are in parallel with the OY axis of the coordinates. The mirror is made of carbon fiber-reinforced polymer, CFRP. It has a total thickness of 7 mm and has ribs for stiffness on the opposite side of the mirror’s reflective face. In order to build the concave mirror, a specific mold was required. The mold was made out of cast iron (Figure 7).

Appl. Sci. 2020, 10, x FOR PEER REVIEW 7 of 18

Appl. Sci. 2020, 10, 7197 7 of 18 Appl. Sci. 2020, 10, x FOR PEER REVIEW 7 of 18

Figure 7. The design (left) and photo (right) of the mold for the forming of the concave mirror.

It features a fine polished concave side according to Equation (1) and with the same dimensions as presented in Figure 5 for the concave mirror. Multiple layers of pre-impregnated carbon fiber with a thickness of 0.3 mm were used for the concave mirror manufacturing. DuringFigureFigure this 7. 7.The process,The design design the (left (left cast) and) and iron photo photo mold (right (right was) of) of the co the moldvered mold for withfor the the formingresin-absorbing forming of of the the concave concave cloth, mirror. mirror.placed in a vacuum bag which was vacuumed at p = 0.01 bar and treated in oven at 180 °C for 6 h. Finally,It features a highly a fine reflectivereflective polished aluminum concave side foil accordin was glued g to on Equation the concave (1) and part part with of of the the concave concavesame dimensions mirror. mirror. Theas presented edges of the in Figure reflectivereflective 5 for surface the concave were mirror. additionally secured with thin strips of aluminum tape, ensuringMultiple the reliable layers contact of pre-impregnated betweenbetween the reflectivereflective carbon fiber foilfoil andand with mirror.mirror. a thickness of 0.3 mm were used for the concave mirror manufacturing. 3.2. The During Design thisof the process, Service Propellant the cast iron Tank mold was covered with resin-absorbing cloth, placed in a vacuum bag which was vacuumed at p = 0.01 bar and treated in oven at 180 °C for 6 h. The service propellant propellant tank tank has has a avolume volume of of 1.6 1.6 L Land and houses houses 462 462 g of g ofN2 Nat2 pattankp =tank 20 =bar.20 The bar. Finally, a highly reflective aluminum foil was glued on the concave part of the concave mirror. Thepressure pressure ptank p=tank 20 bar= 20 is bar the is N the2 pressure N2 pressure inside inside the buffer the bu tankffer tankfor low for temperature low temperature testing. testing. It is also It is The edges of the reflective surface were additionally secured with thin strips of aluminum tape, alsothe pressure the pressure from from which which the tank the tank will will be heated be heated for forthe thehigh high temperature temperature testing. testing. In the In the conditions conditions of ensuring the reliable contact between the reflective foil and mirror. ofthe the experiments experiments conducted conducted at atsea-level sea-level atmospheric atmospheric conditions, conditions, the the propellant propellant tank tank can can reach reach 125 125 °C.◦C. The tank has an outside diameter of 80 mm, an inside diameter of of 50 50 mm mm and and a a length length of of 400 400 mm. mm. 3.2. The Design of the Service Propellant Tank More design details details of of the the service service propellant propellant tank tank ar aree presented presented in inFigure Figure 8. 8It. features It features the thevalve valve and andthe Laval theThe Laval nozzleservice nozzle fitted propellant fitted at the at tank themiddle middle has alength, volume length, wi of withth 1.6an an Ladditional and additional houses two two462 holes holesg of Nfor for2 atthe the ptank pressure = 20 bar. and The temperaturepressure ptank transducer, = 20 bar is which the N 2 are pressure fittedfitted atatinside 120120 ◦the°CC (from(frombuffer thethe tank verticalvertical for low axis)axis) temperature atat eithereither side.testing.side. The It holeis also placedthe pressure vertically from is used which to the mount tank the will solenoid be heated valvevalv fore the to the high propellant temperature tank. testing. The cap In on the theconditions left side of featuresthe experiments a threaded conducted holehole throughthrough at sea-level whichwhich thethe atmospheric pipingpiping carryingcarrying conditions, NN22 is fitted.fitted.the propellant tank can reach 125 °C. The tank has an outside diameter of 80 mm, an inside diameter of 50 mm and a length of 400 mm. More design details of the service propellant tank are presented in Figure 8. It features the valve and the Laval nozzle fitted at the middle length, with an additional two holes for the pressure and temperature transducer, which are fitted at 120 °C (from the vertical axis) at either side. The hole placed vertically is used to mount the solenoid valve to the propellant tank. The cap on the left side features a threaded hole through which the piping carrying N2 is fitted.

Figure 8. D service tank and solenoid valve design.

The service propellant tank was made of 310S-type stainless steel (with a high percentage of Nickel) and features circular stiﬀening ribs. It has a mass of 6100 g, the end caps included, without the valve-nozzle assembly which have the mass of 450 g. The propellant tank is painted with a black mate heat resistant paint. It is placed in the focal axis of the concave mirror, thus absorbingFigure 8. D the service infrared tank light and solenoid emitted valve by the design. infrared lamps and focused by the concave mirror. Appl. Sci. 2020, 10, x FOR PEER REVIEW 8 of 18

The service propellant tank was made of 310S-type stainless steel (with a high percentage of Nickel) and features circular stiffening ribs. It has a mass of 6100 g, the end caps included, without the valve-nozzle assembly which have the mass of 450 g. The propellant tank is painted with a black mate heat resistant paint. It is placed in the focal axis of the concave mirror, thus absorbing the infrared light emitted by the infrared lamps and focused Appl.by the Sci. concave2020, 10, 7197 mirror. 8 of 18

3.3.3.3. TheThe LavalLaval NozzleNozzle TheThe design design of of Laval Laval nozzle nozzle [ 31[31]] and and real real nozzle nozzle can can be be seen seen in in Figure Figure9a,b. 9a,b.

(a) (b)

FigureFigure 9. 9.( a(a)) The The design design of of Laval Laval nozzle; nozzle; and and ( b(b)) Laval Laval nozzle. nozzle.

TheThe nozzle nozzle critical critical section section has has a a diameter diameter of of 3 3 mm. mm. The Thenozzle nozzle is is soldered soldered with with silver silver on on a a fitting fitting whichwhich is is screwed screwed into into the the solenoid solenoid valve valve exhaust exhaust end. end. The The valve valve diameter diameter is is 4 4 mm, mm, less less than than the the fitting fitting innerinner diameter diameter of of 5 5 mm. mm. The The fitting fitting is is screwed screwed into into the the solenoid solenoid valve valve structure structure right right up up to theto the level level of theof the valve. valve. This This design design ensures ensures the the smooth smooth operation operation of the of the valve, valve, with with a minimal a minimal loss loss of N of2. N2. TheThe Laval Laval nozzle nozzle is is made made out out of of 316 316 L L stainless stainless steel steel usingusing aa computercomputer numericalnumerical controlcontrol (CNC)(CNC) millingmilling machine. machine. TheThe upper upper structure structure of of the the hot hot gas gas propulsion propulsion system system consisted consisted of of the the propellant propellant tank’s tank’s supports supports andand ring ring supports. supports. The The lower lower part part of of the the supports supports was was attached attached to the to mirror,the mirror, while while on the on upper the upper part, apart, total a of total three of supporting three supporting rings were rings fixed were on fixed each on support. each support. The supporting The supporting rings were rings laser-cut were fromlaser- 316cut L from stainless 316 steel L sheet.stainless Between steel thesheet. propellant Between tank the support propellant and the tank mirror, support polytetrafluoroethylene and the mirror, (PTFE)polytetrafluoroethylene spacers 5 mm-thick (PTFE) were spacers added 5 to mm-thick prevent thewere concentrated added to prevent heat transfer the concentrated on the mirror. heat Thetransfer ring supportson the mirror. were fixedThe ring with supports the propellant were tankfixed supportswith the usingpropellant M5 screws, tank supports nuts and using washers. M5 Thescrews, propellant nuts and tank washers. was supported The propellant by the threetank was supporting supported rings. by Thethe thre lowere supporting support system rings. was The attachedlower support to the underside system was of the attached mirror withto the the underside screws that of passed the mirror through with the the upper screws support thatsystem, passed Teflonthrough spacers the upper and mirror. support The system, lower Teflon support spacers system and consisted mirror. ofThe a central lower “C”support shaped system support consisted and twoof a lower central support “C” shaped rings forsupport the main and feed two tank. lower The suppo centralrt rings C-shaped for the support main feed had tank. the mirror The central fitted on C- theshaped upper support part and had its the lower mirror part fitted was fitted on the on upper the balance part and arm its end. lower The part central was supportfitted on system the balance was thearm only end. element The central which support connected system the was mirror the andonly itselement upper which elements connected to the balance the mirror arm. and The its lowerupper supportelements rings to the were balance laser arm. cut from The alower 316 L support stainless rings steel were sheet. laser The cut main from propellant a 316 L stainless tank was steel inserted sheet. intoThe the main two propellant rings and tank central was support inserted system. into the The two support rings ringand whichcentral fixes support the rear system. of the The main support tank has a balance pan attached to it. It was used to put weights on it in order to cancel out the imbalance due to weight of the piping, pressure regulator, safety valve and electromagnetic valve mounted on the front part of the system, as shown in Figure 10. Appl. Sci. 2020, 10, x FOR PEER REVIEW 9 of 18 ring which fixes the rear of the main tank has a balance pan attached to it. It was used to put weights on it in order to cancel out the imbalance due to weight of the piping, pressure regulator, safety valve and electromagnetic valve mounted on the front part of the system, as shown in Figure 10. Appl. Sci.When2020 ,the10, 7197infrared lamps were activated, they radiated parallel rays of infrared light to9 ofthe 18 concave mirror which concentrated them on its focal line where the service tank was placed.

Figure 10. TestingTesting equipment during experiments.

InWhen order the to infrared avoid force lamps transducer were activated, damage they during radiated the experiment, parallel rays the of infrared balancelight lever to position the concave was limitedmirror whichby a position concentrated limiter them (Figure on its4), focalwhich line was where fixed the on service the base tank plate. was The placed. position limiter does not allowIn order the totilting avoid of force the balance transducer lever damage beyond during a certain the value. experiment, The travel the balance limit of lever the balance position lever was islimited controlled by a positionby a thin limiter pitch screw (Figure which4), which pushes was on fixed the balance on the basearm plate.on its upper The position side. limiter does not allowThe force the tilting transducer of the balance(Figure lever4) was beyond a high aprecision certain value. one, measuring The travel limita maximum of the balance force of lever 50 N is withcontrolled a precision by a thin of ±0.01 pitch N. screw The which force transducer pushes on thealso balance had the arm option on its to upper record side. 10 readings and the PC liveThe monitoring force transducer of measurements. (Figure4) was a high precision one, measuring a maximum force of 50 N withThe a precision force transducer of 0.01 N. was The mounted force transducer between also the had base the plate option and to the record balance 10 readings lever (Figure and the 4). PC It ± featuredlive monitoring a guide of path measurements. machined into the base plate, which allowed for the transducer to slide across the lengthThe force of the transducer balance arm. was The mounted guide betweenchannel had the basegraduations plate and of the1 mm balance across lever its length. (Figure The4). centerIt featured of the a guideforce transducer path machined was intomarked, the base thus plate, enabli whichng for allowed the easy for read the of transducer the length to of slide the arm across at whichthe length the offorce the was balance applied arm. by The the guide hot gas channel propulsion had graduations system. of 1 mm across its length. The center of theIn force Figure transducer 10, one can was see marked, that the thus pyranometer enabling for is theplaced easy at read a distance of the lengthof 1 m of from the armthe radiation at which sourcesthe force and was at applied 200 mm by theabove hot the gas focal propulsion axis of system. the mirror. It measures the temperature (°C) and irradianceIn Figure (W) 10emitted, one can by seethe thatradiation the pyranometer sources when is placed the arm at is a distancebeing hinged of 1 m into from the the path radiation of the infraredsources andlight at from 200 the mm sources. above the focal axis of the mirror. It measures the temperature (◦C) and irradianceTo further (W) measure emitted byhow the much radiation of the sources infrared when light thereaches arm the is being mirror hinged and how into much the path the mirror of the reflectsinfrared it light back, from the three the sources. thermocouples for the heat dissipation measurement along the distance from the radiationTo further source measure to the how mirror much are of fitted the infrared at a height light ofreaches 1.3, 1.6, theand mirror 1.9 m from and how the radiation much the source. mirror Thereflects elements it back, exposed the three to thermocouples the infrared light for the were heat covered dissipation in gold-plated measurement foil: along the radiation the distance sources from the radiation source to the mirror are fitted at a height of 1.3, 1.6, and 1.9 m from the radiation source. The elements exposed to the infrared light were covered in gold-plated foil: the radiation sources Appl. Sci. 2020, 10, 7197 10 of 18 support, the pyranometer’s hinged arm, the three thermocouples supports and cables, and the cable from the force transducer.

4. Experiments and Results The functionality and principle check of the experimental system was carried out through experiments in two stages, part A and part B. Part A. The first stage in the test campaign comprised performing the tests without the energy source (radiation) concentrated by the mirror cold testing in which the system was a cold gas propulsion system. Part B. The second stage comprised the hot gas testing in which the radiation source which simulates the sun was activated and its energy was concentrated with the mirror. Due to the fact that the measurements were done in atmosphere, the maximum real temperature of the service tank which was achieved was 125 ◦C, the maximum allowed for the operation of the solenoid valve. For this reason, the real case was only simulated for the temperature of 450 ◦C. The temperature of 450 ◦C is the maximum temperature that can be obtained using the available surface of the mirror, considering that the sun rays are uniformly concentrated on the actual surface of the buffer tank. According to the mentioned hypothesis, the absence of radiation is added, so that by applying the Stefan–Boltzmann law, the mentioned value is obtained: r 4 4 I Smirror ∅ = εσT => T = = 469 ◦C (2) σ Stank where ∅ represents the radiant flux, I = 1361 W/m2—solar irradiance, Stefan-Boltsmann constant ( 8) 2 ( 4) 2 σ = 5.67 10 − W/m K − , ε—emissivity, mirror surface Smirror = 0.2527 m , buffer tank surface × 2 Stank = 0.1465 m (corresponding to Φ = 80 mm and a length of 400 mm). Decreasing the tank surface 2 area (to Φ = 20 mm, hence a total surface of 0.0253 m ) will lead to an ideal temperature of 696 ◦C (without radiation losses). Radiation losses q result according to the heat radiation law:

4 4 q = εσ T Tc S (3) h − tank where Th—hot body absolute temperature, and Tc—cold surroundings absolute temperature.

4.1. Part A—Cold Gas Propulsion The testing and validation of the principle of solar–thermal cold gas propulsion were carried out in logical stages, in order to ultimately obtain a conclusive result. It was noticed from the first tests that the opening time of the solenoid valve was a crucial factor for the obtained value of the mean force obtained in a given ∆t and for the total impulse obtained, utilizing a well determined N2 quantity. In this regard, multiple experiments were carried out at low temperature for which the impulses for more opening times of the valve were measured. A fixed quantity of N2 was used (which corresponds to the fill up of the buffer tank at 20 bar). The measured impulses were later summed up in order to make a quantitative comparison. The results for such a test are presented in Table1. For the solenoid valve opening times of 0.1 and 0.3 s, the values were analyzed below. The initial data revealed the fact that there was an error for the force measurement, therefore the calibration of the force transducer was set to start at the value of 10 N and not 0 N. In this regard, it was noticed that there was no major difference between the total impulse obtained with the different opening times. Therefore, the tests were oriented at small valve opening times such that the average force obtained would be in the value vicinity of 1 N, common force values for satellite attitude correction. As it can be noticed in Figure11, an exponential drop of force was obtained when using cold gas propulsion. Appl. Sci. 2020, 10, x FOR PEER REVIEW 11 of 18

Table 1. Total impulse at cold temperature, at several valve opening times.

Opening Time (s) 0.2 0.4 0.5 0.6 Average Average Average Average Isolated Isolated Isolated Isolated Force per Force per Force per Force per Impulse Impulse Impulse Impulse Appl. Sci. 2020, 10, 7197 Impulse Impulse Impulse Impulse11 of 18 (Ns) (Ns) (Ns) (Ns) (N) (N) (N) (N) 2.072 2.309 3.819 2.546 4.351 2.901 4.75 2.969 Table 1. Total impulse at cold temperature, at several valve opening times. 1.555 1.728 2.104 1.365 1.837 1.67 1.654 1.504 1.163 1.292 1.293 1.288Opening Time0.672 (s) 0.747 0.828 0.753 0.721 0.20.811 0.641 0.40.712 0.402 0.50.447 0.257 0.6 0.321 0.487 0.541 0.347 0.434 Average Average Average Average Isolated0.483 0.604 Isolated0.141 0.176 Isolated Isolated Force per Force per Force per Force per Impulse Impulse Impulse Impulse 0.333 Impulse0.416 Impulse Impulse Impulse (Ns) (Ns) (Ns) (Ns) 0.242 0.303(N) (N) (N) (N) 0.196 0.28 2.072 2.309 3.819 2.546 4.351 2.901 4.75 2.969 0.14 0.2 1.555 1.728 2.104 1.365 1.837 1.67 1.654 1.504 0.071 0.101 Total 1.163 1.292 1.293 1.288 0.672 0.747 0.828 0.753 Impulse 0.721 0.811 0.641 0.712 0.402 0.447 0.257 0.321 7.463 8.345 7.262 7.489 Value 0.487 0.541 0.347 0.434 (Ns) 0.483 0.604 0.141 0.176 0.333 0.416 For the solenoid valve opening times of 0.1 and 0.3 s, the values were analyzed below. The initial 0.242data revealed 0.303 the fact that there was an error for the force measurement, therefore the calibration of the0.196 force transducer 0.28 was set to start at the value of 10 N and not 0 N. In this regard,0.14 it was 0.2noticed that there was no major difference between the total impulse obtained with the0.071 different 0.101 opening times. Therefore, the tests were oriented at small valve opening timesTotal such that the average force obtained would be in the value vicinity of 1 N, common force values forImpulse satellite attitude correction. 7.463 8.345 7.262 7.489 ValueAs it can be noticed in Figure 11, an exponential drop of force was obtained when using cold gas propulsion.(Ns)

Figure 11. Force variation with time at the valve opening time of 0.1 at a 10 s interval.

Because the system implies the occurrence of sudden variations (impulses—mechanical in the solenoid valve and gasodynamic in the atmosphere), an oscillatory movement of low amplitude Appl. Sci. 2020, 10, 7197 12 of 18 occurs, but noticeable by the data acquisition system. This oscillatory movement tones down, corresponding to the 10–15 s interval between the opening and closing of the valve. Moreover, it can befrom observed the graphical from the representation graphical representation that the mass that of discharged the mass of ﬂuid discharged from the fluid system from can the be system determined can befrom determined the diﬀerence from the of weight difference (at the of weight beginning (at the and beginning at the end and of the at experiments).the end of the experiments).

4.2.4.2. Part Part B—Hot B—Hot Gas Gas Propulsion Propulsion

InIn order order to to test thethe functionalityfunctionality of of the the test test equipment equipment and and the the principle, principle, a test a attest 80 ◦atC 80 was °C carried was carriedout, the out, temperature the temperature being within being thewithin operation the operation limits of limits the valve of the and valve of the and entire of the test entire equipment. test equipment.After this test, After a comparisonthis test, a comparison between the between impulses the at impulses cold gas at and cold at hotgas gasand was at hot made. gas was made. TheThe illustration illustration of of the the impulse impulse of of the the cold cold gas gas (from (from Figure Figure 11) 11 )in in parallel parallel with with the the impulse impulse from from thethe hot hot gas gas are are represented represented in in Figure Figure 12. 12 .The The tests tests at at low low temperature temperature are are displayed displayed in in blue blue and and the the oneone at at high high temperature, temperature, 80 80 °C,◦C, in in red. red.

FigureFigure 12. 12. ComparisonComparison between between the the impulse impulse for for cold cold gas gas ( (TT00 == 2626 °C)◦C) and and hot hot gas gas (T (Tf f== 8080 °C)◦C) for for the the valvevalve opening opening time time of of 0.1 0.1 s sat at a a10 10 s sinterval. interval.

ItIt can can be be noticed noticed that that the the force force value value for for the the hot hot gas gas is is somewhat somewhat bigger bigger than than the the force force value value obtainedobtained for for the the cold cold gas, gas, in in total, total, as as a asignifican significantt increase increase in in total total impulse impulse is is recorded. recorded. The The impulse impulse increaseincrease is is inin absoluteabsolute values values from from 7.87 7.87 to 8.02to 8. Ns,02 withNs, anwith increase an increase of 2%. Thisof 2%. diff erenceThis difference is percentage is percentagewise so small wise because so small at because small values at small of values opening of time,opening there time, is a there strong is a influence strong influence coming fromcoming the fromvalve the recoil. valve Therefore, recoil. Therefore, at 0.05–0.1 at s,0.05–0.1 a force s, of a small force intensityof small canintensity be noticed can be which noticed acts which in opposition acts in oppositionto the movement, to the movement, but which but destroys which the destroys force distribution. the force distribution. This influence This is influence present utilizing is present the utilizingdetail from the detail Figure from 13, correspondingFigure 13, corresponding to the first to impulse the first from impulse the cold from gas the from cold Figuregas from 11 .Figure 11. TheThe recoil recoil effect effect is is represented represented in in Figures Figures 13 13 an andd 14b. 14b. The The recoil recoil effect effect reduces reduces the the thrust thrust force force onon average average by by 0.3 0.3 N, N, irrespective irrespective of of the the valve valve opening opening time. time. This This is is due due to to the the direction direction of of the the valve valve movement,movement, in in opposition opposition to the directiondirection of of the the thrust thrust force. force. In In Figures Figures 13 13and and 14b, 14b, the the force force would would have havebeen been represented represented by a spike by a inspike the line, in the in the line, opposite in the directionopposite from direction the spike from which the spike is represented which is in representedboth figures. in Therefore,both figures. the Therefore, developed the force developed of the discharged force of the hot discharged gas would hot have gas been would higher have by been0.3 N.higher by 0.3 N. If the case with a valve opening of 0.3 s is analyzed, the valve influence is less pronounced, therefore, there is an increase of 2.4% for the same temperature interval. As an evolution, it was observed in Figure 14a that in this case, there is also a difference between the force and the impulse for the cold gas in comparison with the hot gas. The influence of the valve recoil on the force/impulse can be seen in Figure 14b. The fact that the gas warming inside the tank was carried out with respect to the state Equation must be noticed. Therefore, in Figure 15, the pressure variation with temperature increase xis presented. This variation is displayed in Figure 13 for this case, together with the relevant detail in the warming Appl. Sci. 2020, 10, x FOR PEER REVIEW 13 of 18

Appl. Sci. 2020, 10, 7197 13 of 18

interval. The warming interval is 26–80 ◦C. The decrease in pressure after 80 ◦C corresponds to the valveAppl. Sci. opening 2020, 10 in, x orderFOR PEER toproduce REVIEW the impulses. 13 of 18

Figure 13. The influence of the valve recoil on the impulse.

If the case with a valve opening of 0.3 s is analyzed, the valve influence is less pronounced, therefore, there is an increase of 2.4% for the same temperature interval. As an evolution, it was observed in Figure 14a that in this case, there is also a difference between the force and the impulse for the cold gas in comparison with the hot gas. The influence of the valve recoil on the force/impulseFigureFigure can13. be13. The seenThe inﬂuence influence in Figure of of the14b. the valve valve recoil recoil on on the the impulse. impulse.

(a) (b)

Figure 14. (a) Impulses for the hot and cold gas for Δt = 0.3 s and warming from 26 °C to 80 °C; and Figure 14. (a) Impulses for the hot and cold gas for ∆t = 0.3 s and warming from 26 ◦C to 80 ◦C; and (b) (coldb) cold gas gas force force for for∆t =Δt0.3 = 0.3 fs andfs and valve valve recoil. recoil.

TheThe fact following that the fact gas must warming not be inside neglected: the tank in was the momentcarried out of with a N2 respectquantity to discharge the state Equation from the mustservice be tank, noticed. the temperatureTherefore, in drops( aFigure) suddenly 15, the (aroundpressure 3 variation◦C) because with of thetemperature mass(b) ejection increase outside xis presented.the system. This variation is displayed in Figure 13 for this case, together with the relevant detail in the warmingTheFigure fluid 14. interval. ( temperaturea) Impulses The for warming in the the hot system and interval cold tends gas is for26–80 to Δ increaset = 0.3°C .s Theand because warmingdecrease of from thein pressure tank26 °C thermalto 80 after °C; and inertia,80 °C (b) cold gas force for Δt = 0.3 fs and valve recoil. correspondsthe nitrogen warmingto the valve being opening more in pronounced order to produce as more the quantity impulses. of fluid from the tank is reduced with eachThe valve following opening. fact Furthermore, must not be as neglected: the quantity in the reduces moment with of eacha N2 dischargequantity discharge in the atmosphere, from the The fact that the gas warming inside the tank was carried out with respect to the state Equation servicethe temperature tank, the insidetemperature the tank drops does suddenly not decrease (around as much 3 °C) as because at in case of ofthe the mass first ejection discharge. outside This the can must be noticed. Therefore, in Figure 15, the pressure variation with temperature increase xis system.be noticed in the 75–85 C temperature interval in Figure 15, where at 80 C the valve opens. presented. This variation◦ is displayed in Figure 13 for this case, together◦ with the relevant detail in TheBecause fluid the temperature correction maneuversin the system on thetends orbit to requireincrease (usually) because equal of the impulses, tank thermal the comparison inertia, the of the warming interval. The warming interval is 26–80 °C. The decrease in pressure after 80 °C nitrogenthe cold gaswarming impulses being with more the impulses pronounced after theas more service quantity tank heat-up of fluid from from 22 Cthe to tank60 C–100 is reducedC–125 withC corresponds to the valve opening in order to produce the impulses. ◦ ◦ ◦ ◦ eachwas carriedvalve opening. out. The Furthermore, upper temperature as the quantity limitation reduces was due with to theeach solenoid discharge valve in the material atmosphere, sensitivity, the The following fact must not be neglected: in the moment of a N2 quantity discharge from the which was 125 C. As regards the experiment, of special interest was the total impulse obtained at service tank, the◦ temperature drops suddenly (around 3 °C) because of the mass ejection outside the system. The fluid temperature in the system tends to increase because of the tank thermal inertia, the nitrogen warming being more pronounced as more quantity of fluid from the tank is reduced with each valve opening. Furthermore, as the quantity reduces with each discharge in the atmosphere, the Appl. Sci. 2020, 10, 7197 14 of 18 different initial temperatures of the buffer tank, at a valve opening time of 0.4 s. This opening time was Appl. Sci. 2020, 10, x FOR PEER REVIEW 14 of 18 determined as optimal for a 50 bar total pressure drop from the main tank (which corresponds to a 114temperature g discharged inside mass). the tank By integratingdoes not decrease the force as much related as to at time, in case in of the the necessary first discharge. interval This of 50can bar, be anoticed total impulse in the 75–85 was obtained °C temperature in accordance interval with in Figure Table2 .15, where at 80 °C the valve opens.

FigureFigure 15.15. PressurePressure variationvariation withwith thethe temperaturetemperature warmingwarming fromfrom 2626◦ °CC toto 8080◦ °CC (experiment(experiment fromfrom FigureFigure 12 12).).

BecauseTable the 2. correctionTotal impulse maneuvers value at ∆ont = the0.4 orbit s and re diquirefferent (usually) temperatures equal of theimpulses, buffer tank. the comparison of the cold gas impulses with the impulses after the service tank heat-up from 22 °C to 60 °C–100 °C– Temperature (◦C) Total Impulse (Ns) Related Impulse (Ns/bar) 125 °C was carried out. The upper temperature limitation was due to the solenoid valve material 22 63.559 1.27118 sensitivity, which was 125 °C. As regards the experiment, of special interest was the total impulse 60 70.37 1.4074 obtained at different initial temperatures of the buffer tank, at a valve opening time of 0.4 s. This opening time was determined100 as optimal for 83.643 a 50 bar total pressure 1.67286 drop from the main tank (which corresponds to a 114 g discharged125 mass). By 89.105 integrating the force 1.7821 related to time, in the necessary interval of 50 bar, a total impulse was obtained in accordance with Table 2. Through extrapolating these results for the entire temperature interval, using the functions from Figure 16, aTable significant 2. Total percentage impulse value increase at Δt = (Table 0.4 s and3) can different be obtained. temperatures Two lawsof the of buffer extrapolation tank. were considered, one linearTemperature and the (°C) other one Total polynomial Impulse (Ns) of second Related degree. Impulse (Ns/bar) Usually, the pressure22 and temperature63.559 vary proportionally, with 1.27118 respect to the ideal gas law, by taking into consideration60 no gas loss from70.37 the tank. The experimental 1.4074 data present a semi-linear pressure evolution with100 respect to the temperature83.643 in the 25–125 ◦C interval. 1.67286 The evolution will not be 100% linear given the fact125 that heat will be lost89.105 in space through radiation. 1.7821 The radiation loss is not linear, it varies with T4, where T is the tank temperature. If the radiation is not taken into account, the evolutionThrough is extrapolating linear. If it is these desired results to take for the entire radiation temperature into consideration, interval, using it must the befunctions determined from howFigure the 16, p-T a significant variation drifts percentage away fromincrease the (Table curve. 3) A can balancing be obtained. between Two the laws incoming of extrapolation radiation were flux fromconsidered, the sun, one the linear radiation and giventhe other to the one tank polynomial and the heat of second lost through degree. the radiation must be made. The data extrapolation is presented in Figure 16.

It can be noticed that two extrapolation types were taken into account: a linear extrapolation, which is closer to the intuited value from the start of the project (a 2.5 increase in the total impulse at × maximum temperature) and a polynomial extrapolation. It is impossible to specify an exact value due to the experimental conditions given by the solenoid valve. The fact that a potential 2.7 growth can × Appl. Sci. 2020, 10, x FOR PEER REVIEW 15 of 18

Table 3. Total impulse value at Δt = 0.4 s and different temperatures of the buffer tank (extrapolation).

Impulse Related Percentage Impulse Related Percentage Temperature (Poly) Impulse Growth (Linear) Impulse Growth (°C) (Ns) (Ns/bar) (Polynomial) (Ns) (Ns/bar) (Linear) Tested values 25 63.92 1.27836 0.00 63.31 1.26623 0.00 50 68.14 1.36276 6.60 69.76 1.39518 10.18 75 74.86 1.49716 17.12 76.21 1.52413 20.37 100 84.08 1.68156 31.54 82.65 1.65308 30.55 125 95.80 1.91596 49.88 89.10 1.78203 40.74 Predicted values 150 110.02 2.20036 72.12 95.55 1.91098 50.92 175 126.74 2.53476 98.28 102.00 2.03993 61.10 200 145.96 2.91916 128.35 108.44 2.16888 71.29 225 167.68 3.35356 162.33 114.89 2.29783 81.47 250 191.90 3.83796 200.23 121.34 2.42678 91.65 275 218.62 4.37236 242.03 127.79 2.55573 101.84 300 247.84 4.95676 287.74 134.23 2.68468 112.02 325 279.56 5.59116 337.37 140.68 2.81363 122.21 350 313.78 6.27556 390.91 147.13 2.94258 132.39 375 350.50 7.00996 448.36 153.58 3.07153 142.57 400 389.72 7.79436 509.72 160.02 3.20048 152.76 425 431.44 8.62876 574.99 166.47 3.32943 162.94 450 475.66 9.51316 644.17 172.92 3.45838 173.12

Usually, the pressure and temperature vary proportionally, with respect to the ideal gas law, by taking into consideration no gas loss from the tank. The experimental data present a semi-linear pressure evolution with respect to the temperature in the 25–125 °C interval. The evolution will not

Appl.be 100% Sci. 2020 linear, 10, 7197 given the fact that heat will be lost in space through radiation. The radiation 15loss of 18is not linear, it varies with T4, where T is the tank temperature. If the radiation is not taken into account, the evolution is linear. If it is desired to take the radiation into consideration, it must be determined behow obtained the p-T at variation a temperature drifts ofaway 450 from◦C (temperature the curve. A which balancing can be between reached the at Earth’sincoming low radiation orbit) is flux not tofrom be neglected.the sun, the The radiation polynomial given extrapolation to the tank and is morethe heat optimistic lost through (it approximates the radiation a must 640% be increase), made. althoughThe moredata extrapolation precise results is canpresented be obtained in Figure using 16. a more dedicated valve.

FigureFigure 16.16. Experimentally obtainedobtained impulseimpulse usingusing NN22 at 5050 barbar andand extrapolationextrapolation fromfrom 125125◦ °CC toto 450450◦ °CC atat ∆Δtt == 0.4 s and didifferentﬀerent bubufferﬀer tanktank temperatures.temperatures.

Table 3. Total impulse value at ∆t = 0.4 s and diﬀerent temperatures of the buﬀer tank (extrapolation). It can be noticed that two extrapolation types were taken into account: a linear extrapolation, which is closer to theImpulse intuited valuRelatede from thePercentage start of the projectImpulse (a 2.5 x Relatedincrease inPercentage the total impulse Temperature at maximum temperature)(Poly) and a Impulsepolynomial extrapolation.Growth (Linear) It is impossibleImpulse to specifyGrowth an exact value ( C) ◦ (Ns) (Ns/bar) (Polynomial) (Ns) (Ns/bar) (Linear) Tested values 25 63.92 1.27836 0.00 63.31 1.26623 0.00 50 68.14 1.36276 6.60 69.76 1.39518 10.18 75 74.86 1.49716 17.12 76.21 1.52413 20.37 100 84.08 1.68156 31.54 82.65 1.65308 30.55 125 95.80 1.91596 49.88 89.10 1.78203 40.74 Predicted values 150 110.02 2.20036 72.12 95.55 1.91098 50.92 175 126.74 2.53476 98.28 102.00 2.03993 61.10 200 145.96 2.91916 128.35 108.44 2.16888 71.29 225 167.68 3.35356 162.33 114.89 2.29783 81.47 250 191.90 3.83796 200.23 121.34 2.42678 91.65 275 218.62 4.37236 242.03 127.79 2.55573 101.84 300 247.84 4.95676 287.74 134.23 2.68468 112.02 325 279.56 5.59116 337.37 140.68 2.81363 122.21 350 313.78 6.27556 390.91 147.13 2.94258 132.39 375 350.50 7.00996 448.36 153.58 3.07153 142.57 400 389.72 7.79436 509.72 160.02 3.20048 152.76 425 431.44 8.62876 574.99 166.47 3.32943 162.94 450 475.66 9.51316 644.17 172.92 3.45838 173.12

Given the mission of the satellite, the propellant mass onboard can vary from satellite to satellite. As a comparison value, the propellant mass is 1/5 the mass of the nano-satellite Adelis-SAMSON, Appl. Sci. 2020, 10, 7197 16 of 18 which is fitted with a cold gas propulsion system. Adelis-SAMSON carries 468 gr of Krypton at 160 bar and 30 ◦C. This ensures a total impulse of >150 Ns [32]. In comparison, the hot gas propulsion system develops a total impulse of 172 Ns with nitrogen at 450 ◦C and 27 bar, heated from 25 ◦C and 20 bar. Despite the fact that direct comparisons are difficult, as the tank volume and pressure are different for both solutions, it can be noticed that the hot gas propulsion system offers a 14.6% increase in the total impulse, with same mass of propellant but at six times-lower propellant pressure.

5. Conclusions The conclusion of the tests and the subsequent data extrapolations is that the specific impulse of the hot gas propulsion system increases by 173.24% over the obtained specific impulse at 25 ◦C where no heating of the tank was carried out. The temperature of 450 ◦C is the temperature which the propellant tank reaches in space, when its heating is carried out by the mirror. Therefore, the operational life of an LEO satellite increases 2.73 times. The maximum thrust force is 8 N when the valve opening time is 0.4 s. The maximum thrust force which the hot gas propulsion system achieves is 9.21 N, at 27 bar, achieved after heating up the buffer tank to 125 ◦C and opening the valve for 0.3 s. The limitations of this study are measurement related. The discharged gas speed at the nozzle was neither measured or determined. Therefore, the specific impulse could not be calculated and as a result, the total impulse (which is correlated with the autonomy of the hot gas propulsion system) was not determined from the specific impulse. By determining the total impulse from the specific impulse, less result errors would have occurred and the calculation procedure would have been more straightforward. Therefore, for future experiments, a lightweight propellant tank will be developed with a 0.7 L volume and fitted with six valves and nozzles In the future, it is of interest to test the new concept for a wider range of temperatures, up to 450 ◦C, which is the mirror temperature threshold. It is intended to carry out the tests in vacuum and temperature conditions similarly to the values from the cosmic space for LEO. Important aspects which will arise when deploying the LEO satellite onto the orbit are control related. The linear parabolic mirror destined for space will comprise two symmetrical parts which will be folded until the first correction maneuver. When they unfold, they will create the complete mirror. The mirror will be silver plated and the valve-nozzle assemblies will be fitted in a manner to cover all ranges of necessary correction movements, which means a 360◦ action. Problems that will arise are control-related: The correction maneuver can alter the position of the satellite in space, therefore, all corrections will be limited to a valve opening time of maximum 0.3 s. Therefore, no gas will be wasted. The mirror destined for testing has a mass of 2105 g. The mirror destined for the space application will have a mass of 1100 g and will be made out of 2 mm-thick CFRP. Furthermore, the service tank/buffer tank will have a mass of 460 g. Including all the valves, piping and sensors, the total mass of the system destined for space will be of 3150 g, therefore 1150 g heavier than a comparable cold gas propulsion system, in terms of tank capacity, quantity of propellant carried onboard and developed thrust.

Author Contributions: Conceptualization, C.S. and G.C.; methodology, H.S, .; software, A.R. and A.T.; validation,

A.T.; writing—original draft preparation, H.S, ., C.S., G.C., A.T., T.T., A.R.; writing—review and editing, H.S, ., G.C.; supervision V.S.; project administration, V.S. All authors have read and agreed to the published version of the manuscript. Funding: This work was carried out within STAR, “STRAUSS” Project, Grant no. 130/2017, Project supported by the Romanian Minister of Research and Innovation. Conﬂicts of Interest: The authors declare no conﬂict of interest. Appl. Sci. 2020, 10, 7197 17 of 18

References

1. ESA Space Debris Office, 2020, Space Debris by the Numbers. Available online: https://www.esa.int/Safety_ Security/Space_Debris/Space_debris_by_the_numbers (accessed on 10 May 2020). 2. Adushkin, V.; Aksenov, O.Y.; Veniaminov, S.; Kozlov, S.; Tyurenkova, V. The small orbital debris population and its impact on space activities and ecological safety. Acta Astronaut. 2020.[CrossRef] 3. Sutton, G.P.; Biblarz, O. Rocket Propulsion Elements, 7th ed.; John Wiley & Sons: New York, NY, USA, 2001; pp. 679–688. 4. Rovey, J.L.; Lyne, C.T.; Mundahl, A.J.; Rasmont, N.; Glascock, M.S.; Wainwright, M.J.; Berg, S.P. Review of multimode space propulsion. Prog. Aerosp. Sci. 2020, 118, 100627. [CrossRef] 5. Silva, M.A.; Guerrieri, D.C.; Cervone, A.; Gill, E. A review of MEMS micropropulsion technologies for CubeSats and PocketQubes. Acta Astronaut. 2018, 143, 234–243. [CrossRef] 6. Johnson, L.; Whorton, M.; Heaton, A.; Pinson, R.; Laue, G.; Adams, C. NanoSail-D: A solar sail demonstration mission. Acta Astronaut. 2011, 68, 571–575. [CrossRef] 7. Lappas, V.; Adeli, N.; Visagie, L.; Fernandez, J.; Theodorou, T.; Steyn, W.; Perren, M. CubeSail: A low cost CubeSat based solar sail demonstration mission. Adv. Space Res. 2011, 48, 1890–1901. [CrossRef] 8. Kennedy, F.G. Solar Thermal Propulsion for Microsatellite Manoeuvring; University of Surrey: Guildford, UK, 2004. 9. John, E.; Canady, J.E., Jr.; John, L.; Allen, J.L., Jr. Illumination from Space with Orbiting Solar-Reflector Spacecraft; NASA Technical Paper 2065; NASA: Washington, DC, USA, 1982. 10. Li, X.; Duan, B.; Song, L.; Yang, Y.; Zhang, Y.; Wang, D. A new concept of space solar power satellite. Acta Astronaut. 2017, 136, 182–189. [CrossRef] 11. Lior, N. Mirrors in the sky: Status, sustainability, and some supporting materials experiments. Renew. Sustain. Energy Rev. 2013, 18, 401–415. [CrossRef] 12. Glaser, P.E. Power from the Sun: Its Future. Science 1968, 162, 857–861. [CrossRef][PubMed] 13. Grossman, G.; Williams, G. Inflatable Concentrators for Solar Propulsion and Dynamic Space Power. J. Sol. Energy Eng. 1990, 112, 229–236. [CrossRef] 14. Holmes, M. Solar Rocket Propulsion: Ground and Space Technology Demonstration. In Proceedings of the 12th Annual Advanced Space Propulsion Workshop, Huntsville, AL, USA, 3–5 April 2001. 15. Schenk, M.; Viquerat, A.; Seffen, K.A.; Guest, S.D. Review of Inflatable Booms for Deployable Space Structures: Packing and Rigidization. J. Spacecr. Rocket. 2014, 51, 762–778. [CrossRef] 16. Ehricke, K.A. Space light: Space industrial enhancement of the solar option. Acta Astronaut. 1979, 6, 1515–1633. [CrossRef] 17. Herbert, J. Kramer Observation of the Earth and Its Environment: Survey of Missions and Sensors. Front Cover; Springer Science & Business Media: Berlin/Heidelberg, Germany, 2002. 18. Sistach, L.A. The moon orbital mirror. In Proceedings of the 54th International Astronautical Congress of the International Astronautical Federation, Bremen, Germany, 29 September 2003; The International Academy of Astronautics, and the International Institute of Space Law. 19. Ehricke, K.A. The Solar-Powered Space Ship; General Dynamics/Astronautics: San Diego, CA, USA, 1959. 20. Leverone, F.; Cervone, A.; Gill, E. Cost analysis of solar thermal propulsion systems for microsatellite applications. Acta Astronaut. 2019, 155, 90–110. [CrossRef] 21. Leverone, F.A.; Cervone, M.; Pini, E.; Gill, P. Feasibility of an integrated solar thermal power and propulsion system for small satellites. In Proceedings of the 68th International Astronautical Congress (IAC), Adelaide, Australia, 25–29 September 2017. 22. Assad, A. Remote Sensing Advanced Techniques and Platforms—Chapter 20—Cold Gas Propulsion System—An Ideal Choice for Remote Sensing Small Satellites. IntechOpen 2012. Available online: https://www.intechopen.com/books/remote-sensing-advanced-techniques-and-platforms/cold-gaspropulsion-system-an-ideal-choice-for-remote-sensing-small-satellites (accessed on 15 August 2020). 23. Ghasemi, M.; Rezaeiha, A. Design & Development of a Flight-Model Cold Gas Propulsion System, ISTS2011-a-14. In Proceedings of the 28th International Symposium on Space Science and Technology, Okinawa, Japan, 5–12 June 2011. Appl. Sci. 2020, 10, 7197 18 of 18

24. Mojtaba, G.; Abdolrahim, R. Flight Qualification Tests of a Cold Gas Propulsion System for a Small Satellite. In Proceedings of the AJCPP 2012 Asian Joint Conference on Propulsion and Power, Xi’an, China, 1–4 March 2012. 25. Kindracki, J.; Tur, K.; Paszkiewicz, P.; M˛e˙zyk,Ł.; Boruc, Ł.; Wolański,P. Experimental research on low-cost cold gas propulsion for a space robot platform. Aerosp. Sci. Technol. 2017, 62, 148–157. [CrossRef] 26. Ranjan, R.; Karthikeyan, K.; Riaz, F.; Chou, S. Cold gas propulsion microthruster for feed gas utilization in micro satellites. Appl. Energy 2018, 220, 921–933. [CrossRef] 27. Hristie, J. Transport Processes and Unit Operations, 3rd ed.; P T R Prentice-Hall, Inc.: Englewood Cliffs, NJ, USA, 1993; ISBN 0-13-045253-X. 28. Lean, J.L. The sun’s variable radiation and its relevance for earth. Annu. Rev. Astron. Astrophys. 1997, 35, 33–67. [CrossRef] 29. Available online: https://ventiltechnik.de/fileadmin/gsr/media/Data_sheets/gsr-data-sheet-solenoid-valve- type-52hd.pdf (accessed on 1 September 2020). 30. ¸Serbescu,H.; Sandu, C.; Vintila, I.S.; Radu, A.; Niculescu, F. Development of a parabolic mirror using advanced materials used for an environment friendly propulsion system. In Recent Advances on Environment, Chemical Engineering and Materials, AIP Conference Proceedings; AIP Publishing: College Park, MD, USA, 2018; pp. 020005-1–020005-9. [CrossRef] 31. Brian, J. AA210A Course Reader Fundamentals of Compressible Flow; Department of Aeronautics and Astronautics Stanford University: Stanford, CA, USA, 2018. 32. Zaberchik, M.; Lev, D.R.; Edlerman, E.; Kaidar, A. Fabrication and Testing of the Cold Gas Propulsion System Flight Unit for the Adelis-SAMSON Nano-Satellites. Aerospace 2019, 6, 91. [CrossRef]

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