RN
R. Noomen
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The current state of in-orbit refuelling involves launching propellant from the Earth’s surface in a single use refuelling craft, often to transfer hydrazine to the customer. This method is a logical first step however this architecture is not reusable, and it centres around a toxic and carcinogenic propellant. An architecture is proposed where hydrogen peroxide, a green oxidiser useful in both propulsion and power systems, is created from water ice in the solar system and refuelled with a reusable refuelling craft. First order sizing of the craft is conducted showing the viability of refuelling routes from Deimos, Phobos and the Moon. Prototype testing of a propellant transfer mechanism has shown the promise of using a piston-based transfer system, and the results of testing are used to better estimate the mass of a potential reusable refuelling craft.
...
The current state of in-orbit refuelling involves launching propellant from the Earth’s surface in a single use refuelling craft, often to transfer hydrazine to the customer. This method is a logical first step however this architecture is not reusable, and it centres around a toxic and carcinogenic propellant. An architecture is proposed where hydrogen peroxide, a green oxidiser useful in both propulsion and power systems, is created from water ice in the solar system and refuelled with a reusable refuelling craft. First order sizing of the craft is conducted showing the viability of refuelling routes from Deimos, Phobos and the Moon. Prototype testing of a propellant transfer mechanism has shown the promise of using a piston-based transfer system, and the results of testing are used to better estimate the mass of a potential reusable refuelling craft.
Interplanetary CubeSat missions enable low-cost exploration missions, in which the design of a primary propulsion system becomes crucial. Emphasis is put on an autonomous Mars mission starting from a parking orbit around Earth: chemical propulsion systems are characterized by high thrust levels and allow fast Earth escape manoeuvres. In recent years, new green propellant options have been analysed and are available for CubeSat propulsion.
A preliminary mission analysis is performed considering the performance parameters of state-of-the-art CubeSat chemical propulsion systems. Afterwards, a trade-off leads to the choice of propulsion system type and propellant. The propulsion system challenge is to fit in a CubeSat standardized volume, which can range up to 24U: a final design is produced, showing the feasibility of the mission adopting multiple COTS components. ...
A preliminary mission analysis is performed considering the performance parameters of state-of-the-art CubeSat chemical propulsion systems. Afterwards, a trade-off leads to the choice of propulsion system type and propellant. The propulsion system challenge is to fit in a CubeSat standardized volume, which can range up to 24U: a final design is produced, showing the feasibility of the mission adopting multiple COTS components. ...
Interplanetary CubeSat missions enable low-cost exploration missions, in which the design of a primary propulsion system becomes crucial. Emphasis is put on an autonomous Mars mission starting from a parking orbit around Earth: chemical propulsion systems are characterized by high thrust levels and allow fast Earth escape manoeuvres. In recent years, new green propellant options have been analysed and are available for CubeSat propulsion.
A preliminary mission analysis is performed considering the performance parameters of state-of-the-art CubeSat chemical propulsion systems. Afterwards, a trade-off leads to the choice of propulsion system type and propellant. The propulsion system challenge is to fit in a CubeSat standardized volume, which can range up to 24U: a final design is produced, showing the feasibility of the mission adopting multiple COTS components.
A preliminary mission analysis is performed considering the performance parameters of state-of-the-art CubeSat chemical propulsion systems. Afterwards, a trade-off leads to the choice of propulsion system type and propellant. The propulsion system challenge is to fit in a CubeSat standardized volume, which can range up to 24U: a final design is produced, showing the feasibility of the mission adopting multiple COTS components.
Some propellants currently used in space propulsion have the disadvantage of being dangerous to humans and the environment. A wellknown example of this are hydrazine and its derivatives which are toxic, corrosive, and carcinogenic. These disadvantages have led to a search for less hazardous storable liquid propellants, often referred to as green propellants. A green propellant is safer to use and handle and will therefore bring down the costs related to production, storage and handling. The main reason for the use of hydrazine is its hypergolicity with common oxidizers like NTO and nitric acid. Hypergolicity is the property that a fuel and oxidizer ignite spontaneously when brought in contact without the need for an external source. This property is beneficial since it eliminates the need for an ignition system, thereby making the propulsion system more simple, reliable, and cheaper. It is therefore desired that a replacement propellant also shows this property with common oxidizers.
During this thesis two methods are explored with the goal of creating a green hypergolic propellant combination. One based on catalytically enhanced ethanol and high concentration hydrogen peroxide. The other based on a pyrophoric liquid that is added to ethanol. There have already been efforts made before by adding catalyst particles or a strong reducer to a hydrocarbon fuel like kerosene or ethanol. The problem with using catalyst or strong reducer particles is the difficulty of creating a homogeneous mixture. Due to the liquid nature of the parent fuel, the particles will start to separate from the fuel and sink to the bottom of the container. To overcome this problem an organic gelling agent is added to the fuel to increase its viscosity, thereby increasing the sedimentation time. This will not only increase the shelf life of the fuel but also decrease the vapor pressure, making it less flammable and safer to work with, and reduce storage problems like propellant sloshing and spilling. Increasing the viscosity also makes it more difficult to transfer the fuel through the feed system and achieve proper atomization. However, by applying shear force on the fuel it shows shear thinning behavior, decreasing the viscosity close to that of the parent properties making it easy to use in existing propulsion systems designed for liquids. A second effect of the organic gelling agent is that due to its energetic nature it participates in the combustion. Therefore, the amount of catalyst needed can be reduced to a negligible amount while still achieving good performance properties like ignition delay time. This is verified by means of a drop test.
As a second approach, instead of adding solid catalyst particles a pyrophoric liquid is added to ethanol. A liquid allows for easier mixing and creating of a homogeneous mixture resulting in a longer shelf life compared to using catalyst particles. The pyrophoric liquid also fully participates in combustion resulting in increased performance compared to catalyst particles that do not combust. Due to the reactive nature of the pyrophoric liquid this propellant formulation is expected to be able to achieve hypergolicity with multiple common oxidizers. This results in a versatile system which can be used in current propulsion systems without extensive modification required. By eliminating the need for an ignition system, the propulsion system is simplified and its reliability is increased. ...
During this thesis two methods are explored with the goal of creating a green hypergolic propellant combination. One based on catalytically enhanced ethanol and high concentration hydrogen peroxide. The other based on a pyrophoric liquid that is added to ethanol. There have already been efforts made before by adding catalyst particles or a strong reducer to a hydrocarbon fuel like kerosene or ethanol. The problem with using catalyst or strong reducer particles is the difficulty of creating a homogeneous mixture. Due to the liquid nature of the parent fuel, the particles will start to separate from the fuel and sink to the bottom of the container. To overcome this problem an organic gelling agent is added to the fuel to increase its viscosity, thereby increasing the sedimentation time. This will not only increase the shelf life of the fuel but also decrease the vapor pressure, making it less flammable and safer to work with, and reduce storage problems like propellant sloshing and spilling. Increasing the viscosity also makes it more difficult to transfer the fuel through the feed system and achieve proper atomization. However, by applying shear force on the fuel it shows shear thinning behavior, decreasing the viscosity close to that of the parent properties making it easy to use in existing propulsion systems designed for liquids. A second effect of the organic gelling agent is that due to its energetic nature it participates in the combustion. Therefore, the amount of catalyst needed can be reduced to a negligible amount while still achieving good performance properties like ignition delay time. This is verified by means of a drop test.
As a second approach, instead of adding solid catalyst particles a pyrophoric liquid is added to ethanol. A liquid allows for easier mixing and creating of a homogeneous mixture resulting in a longer shelf life compared to using catalyst particles. The pyrophoric liquid also fully participates in combustion resulting in increased performance compared to catalyst particles that do not combust. Due to the reactive nature of the pyrophoric liquid this propellant formulation is expected to be able to achieve hypergolicity with multiple common oxidizers. This results in a versatile system which can be used in current propulsion systems without extensive modification required. By eliminating the need for an ignition system, the propulsion system is simplified and its reliability is increased. ...
Some propellants currently used in space propulsion have the disadvantage of being dangerous to humans and the environment. A wellknown example of this are hydrazine and its derivatives which are toxic, corrosive, and carcinogenic. These disadvantages have led to a search for less hazardous storable liquid propellants, often referred to as green propellants. A green propellant is safer to use and handle and will therefore bring down the costs related to production, storage and handling. The main reason for the use of hydrazine is its hypergolicity with common oxidizers like NTO and nitric acid. Hypergolicity is the property that a fuel and oxidizer ignite spontaneously when brought in contact without the need for an external source. This property is beneficial since it eliminates the need for an ignition system, thereby making the propulsion system more simple, reliable, and cheaper. It is therefore desired that a replacement propellant also shows this property with common oxidizers.
During this thesis two methods are explored with the goal of creating a green hypergolic propellant combination. One based on catalytically enhanced ethanol and high concentration hydrogen peroxide. The other based on a pyrophoric liquid that is added to ethanol. There have already been efforts made before by adding catalyst particles or a strong reducer to a hydrocarbon fuel like kerosene or ethanol. The problem with using catalyst or strong reducer particles is the difficulty of creating a homogeneous mixture. Due to the liquid nature of the parent fuel, the particles will start to separate from the fuel and sink to the bottom of the container. To overcome this problem an organic gelling agent is added to the fuel to increase its viscosity, thereby increasing the sedimentation time. This will not only increase the shelf life of the fuel but also decrease the vapor pressure, making it less flammable and safer to work with, and reduce storage problems like propellant sloshing and spilling. Increasing the viscosity also makes it more difficult to transfer the fuel through the feed system and achieve proper atomization. However, by applying shear force on the fuel it shows shear thinning behavior, decreasing the viscosity close to that of the parent properties making it easy to use in existing propulsion systems designed for liquids. A second effect of the organic gelling agent is that due to its energetic nature it participates in the combustion. Therefore, the amount of catalyst needed can be reduced to a negligible amount while still achieving good performance properties like ignition delay time. This is verified by means of a drop test.
As a second approach, instead of adding solid catalyst particles a pyrophoric liquid is added to ethanol. A liquid allows for easier mixing and creating of a homogeneous mixture resulting in a longer shelf life compared to using catalyst particles. The pyrophoric liquid also fully participates in combustion resulting in increased performance compared to catalyst particles that do not combust. Due to the reactive nature of the pyrophoric liquid this propellant formulation is expected to be able to achieve hypergolicity with multiple common oxidizers. This results in a versatile system which can be used in current propulsion systems without extensive modification required. By eliminating the need for an ignition system, the propulsion system is simplified and its reliability is increased.
During this thesis two methods are explored with the goal of creating a green hypergolic propellant combination. One based on catalytically enhanced ethanol and high concentration hydrogen peroxide. The other based on a pyrophoric liquid that is added to ethanol. There have already been efforts made before by adding catalyst particles or a strong reducer to a hydrocarbon fuel like kerosene or ethanol. The problem with using catalyst or strong reducer particles is the difficulty of creating a homogeneous mixture. Due to the liquid nature of the parent fuel, the particles will start to separate from the fuel and sink to the bottom of the container. To overcome this problem an organic gelling agent is added to the fuel to increase its viscosity, thereby increasing the sedimentation time. This will not only increase the shelf life of the fuel but also decrease the vapor pressure, making it less flammable and safer to work with, and reduce storage problems like propellant sloshing and spilling. Increasing the viscosity also makes it more difficult to transfer the fuel through the feed system and achieve proper atomization. However, by applying shear force on the fuel it shows shear thinning behavior, decreasing the viscosity close to that of the parent properties making it easy to use in existing propulsion systems designed for liquids. A second effect of the organic gelling agent is that due to its energetic nature it participates in the combustion. Therefore, the amount of catalyst needed can be reduced to a negligible amount while still achieving good performance properties like ignition delay time. This is verified by means of a drop test.
As a second approach, instead of adding solid catalyst particles a pyrophoric liquid is added to ethanol. A liquid allows for easier mixing and creating of a homogeneous mixture resulting in a longer shelf life compared to using catalyst particles. The pyrophoric liquid also fully participates in combustion resulting in increased performance compared to catalyst particles that do not combust. Due to the reactive nature of the pyrophoric liquid this propellant formulation is expected to be able to achieve hypergolicity with multiple common oxidizers. This results in a versatile system which can be used in current propulsion systems without extensive modification required. By eliminating the need for an ignition system, the propulsion system is simplified and its reliability is increased.
As a part of a Future Launcher Preparatory Programme contract issued by the European Space Agency, the aerospace startup company Dawn Aerospace is developing an integral, additive manufactured (AM), liquid rocket engine. This 2.5 kN class engine is propelled by the storable propellants 90% hydrogen peroxide and kerosene and used in the application of a (sub)orbital spaceplane. In this study, a preliminary design for the thrust chamber of the AM engine is proposed, which is realized from the superalloy Inconel 718. To avoid the chamber wall from melting under the extreme thermal loads which originate from the ongoing combustion reaction, both regenerative cooling and film cooling with hydrogen peroxide are considered. The proposed designs are based on a custom-developed 2-D numerical analysis code, which considers non-linear steady-state heat transfer and linear elastic structural deformation. The results of this code are validated against hot-fire tests of both regenerative cooled and film cooled thrust chambers, propelled by 90% hydrogen peroxide and kerosene. Moreover, an immersion screening test campaign is conducted to validate the chemical compatibility of AM Inconel 718 with 90% hydrogen peroxide. This thesis compares the proposed AM chamber designs to the current bimetallic chamber design that is already developed and hot-fired by Dawn Aerospace. The bimetallic design comprises a coated copper-alloy liner and relies mostly on non-AM production techniques. The coating is introduced to improve the chemical compatibility with hydrogen peroxide. Simulations show that the total available delta-v of the spaceplane is reduced from 3.46 km/s to 3.38 km/s when the bimetallic design is replaced by an AM Inconel 718 design, relying on regenerative cooling and film cooling. On the contrary, the AM design may offer a thrust chamber dry mass reduction of more than 75%, whereas the total number of sealing interfaces can be reduced by more than 50%. In addition, the inspection-heavy coating, which is plated onto the copper alloy liner of the current bimetallic engine design, can be removed in the AM design.
...
As a part of a Future Launcher Preparatory Programme contract issued by the European Space Agency, the aerospace startup company Dawn Aerospace is developing an integral, additive manufactured (AM), liquid rocket engine. This 2.5 kN class engine is propelled by the storable propellants 90% hydrogen peroxide and kerosene and used in the application of a (sub)orbital spaceplane. In this study, a preliminary design for the thrust chamber of the AM engine is proposed, which is realized from the superalloy Inconel 718. To avoid the chamber wall from melting under the extreme thermal loads which originate from the ongoing combustion reaction, both regenerative cooling and film cooling with hydrogen peroxide are considered. The proposed designs are based on a custom-developed 2-D numerical analysis code, which considers non-linear steady-state heat transfer and linear elastic structural deformation. The results of this code are validated against hot-fire tests of both regenerative cooled and film cooled thrust chambers, propelled by 90% hydrogen peroxide and kerosene. Moreover, an immersion screening test campaign is conducted to validate the chemical compatibility of AM Inconel 718 with 90% hydrogen peroxide. This thesis compares the proposed AM chamber designs to the current bimetallic chamber design that is already developed and hot-fired by Dawn Aerospace. The bimetallic design comprises a coated copper-alloy liner and relies mostly on non-AM production techniques. The coating is introduced to improve the chemical compatibility with hydrogen peroxide. Simulations show that the total available delta-v of the spaceplane is reduced from 3.46 km/s to 3.38 km/s when the bimetallic design is replaced by an AM Inconel 718 design, relying on regenerative cooling and film cooling. On the contrary, the AM design may offer a thrust chamber dry mass reduction of more than 75%, whereas the total number of sealing interfaces can be reduced by more than 50%. In addition, the inspection-heavy coating, which is plated onto the copper alloy liner of the current bimetallic engine design, can be removed in the AM design.
An important topic in the field of rocket propulsion is the development of a new green advanced hypergolic rocket propulsion system. The current hypergolic propellants have one large drawback, namely the toxic properties. The injector has a crucial role in the performance of a liquid engine. Therefore, the next step in the development of this novel propulsion system is the development of the injector. The purpose of this master thesis is to design, manufacture, and test this novel injector. The injector is designed to operate in monopropellant and bi-propellant modes, which is supported by the new green hypergolic propellants. In total, four different configurations of the injector are developed. Different testing objectives are defined for the experiments to test all the injector configurations. These testing objectives are atomization behavior, atomization performance, the mixing process in bi-propellant mode operation, and atomized flow in combination with the heating element.
...
An important topic in the field of rocket propulsion is the development of a new green advanced hypergolic rocket propulsion system. The current hypergolic propellants have one large drawback, namely the toxic properties. The injector has a crucial role in the performance of a liquid engine. Therefore, the next step in the development of this novel propulsion system is the development of the injector. The purpose of this master thesis is to design, manufacture, and test this novel injector. The injector is designed to operate in monopropellant and bi-propellant modes, which is supported by the new green hypergolic propellants. In total, four different configurations of the injector are developed. Different testing objectives are defined for the experiments to test all the injector configurations. These testing objectives are atomization behavior, atomization performance, the mixing process in bi-propellant mode operation, and atomized flow in combination with the heating element.
Master thesis
(2021)
-
Y.S. le Grand, B.T.C. Zandbergen, M. Merino, H.J. de Blank, A. Cervone, R. Noomen
Electric propulsion is generally characterized by higher specific impulse than traditional chemical rockets, which gives them propellant saving benefits. These have been applied in deep-space missions and micro-satellites. A Pulsed Plasma Thruster consists of two electrodes around a block of teflon. A pulsed discharge ablates the teflon and ionizes the vapor. This plasma forms a current bridge between the electrodes which is accelerated outwards by the self-induced magnetic field, producing thrust.
The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated. ...
The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated. ...
Electric propulsion is generally characterized by higher specific impulse than traditional chemical rockets, which gives them propellant saving benefits. These have been applied in deep-space missions and micro-satellites. A Pulsed Plasma Thruster consists of two electrodes around a block of teflon. A pulsed discharge ablates the teflon and ionizes the vapor. This plasma forms a current bridge between the electrodes which is accelerated outwards by the self-induced magnetic field, producing thrust.
The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated.
The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated.
The commercialization of the space industry has led to a reduction in size and weight of satellites and launching vehicles, which have effectively reduced the cost of space services. The costs of launching payloads to space can be significantly reduced when the launching vehicles are reusable. The two-stage-to-orbit system with a winged, reusable first stage vehicle, is deemed to offer benefits in terms of operations, as it can take-off and land from a runway.
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.
...
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.
...
The commercialization of the space industry has led to a reduction in size and weight of satellites and launching vehicles, which have effectively reduced the cost of space services. The costs of launching payloads to space can be significantly reduced when the launching vehicles are reusable. The two-stage-to-orbit system with a winged, reusable first stage vehicle, is deemed to offer benefits in terms of operations, as it can take-off and land from a runway.
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.
Potassium Nitrate Sorbitol Propellant
Experimental Investigation of Solid Propellant Characteristics
Master thesis
(2019)
-
Martin Olde, Eberhard Gill, Barry Zandbergen, Ron Noomen, Botchu Vara Siva Jyoti
KNO3-Sugar propellants, also known as rocket candy, form a group of simple, cheap and safe solid propellants that are used extensively in student and amateur rocketry communities. One of the most frequently used compositions is KNO3-sorbitol (KNSB) with a typical 65/35 ratio by mass. This composition is used extensively by Delft Aerospace Rocket Engineering (DARE) for small experimental launches including the record-breaking flight of Stratos I in 2009 to 12.3 km altitude. However, KNSB propellant quality has been inconsistent: propellant density is occasionally below 85% in combination with large surface defects. Besides a very high grain rejection rate this has resulted in several explosive failures of new experimental motors in recent years.
...
KNO3-Sugar propellants, also known as rocket candy, form a group of simple, cheap and safe solid propellants that are used extensively in student and amateur rocketry communities. One of the most frequently used compositions is KNO3-sorbitol (KNSB) with a typical 65/35 ratio by mass. This composition is used extensively by Delft Aerospace Rocket Engineering (DARE) for small experimental launches including the record-breaking flight of Stratos I in 2009 to 12.3 km altitude. However, KNSB propellant quality has been inconsistent: propellant density is occasionally below 85% in combination with large surface defects. Besides a very high grain rejection rate this has resulted in several explosive failures of new experimental motors in recent years.