Rocket Performance Research Articles

A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced.

Effect on the performance of a hybrid rocket using AP in the PVC and HTPB-based hybrid fuel

Hybrid rocket is a type of chemical rocket propulsion system where the fuel and the oxidizer are stored separated and in different state of aggregation. Typically, the fuel is in the solid state and the oxidizer is either in the liquid or gaseous state. It offers various advantages over other type of chemical rockets such as throattability, controllability and reduced complexity. It also has some drawbacks such as low regression rate, poor combustion efficiency, higher sliver losses, etc. Various methods have been adopted to overcome these drawbacks and improve the performance of hybrid rocket. The present study focus on the improvement of regression rate with the addition of oxidizer particles in fuel content. Ammonium Perchlorate is used as an additive in Hydroxyl Terminated Polybutadiene (HTPB) and Poly Vinyl Chloride (PVC) based fuel grain. From the experimentations, it was observed that the thrust, pressure and regression rate of the HTPB fuel grain with additive have higher improvement than with the PVC based fuel grain. The sliver losses are also reduced due to uniform burning after addition of oxidizer in fuel grain. The overall study strongly suggests that addition of oxidizer in limited quantity can be used to overcome the drawbacks of hybrid propellant rocket.

Development of hydrogen peroxide based micro thruster utilizing biomass-based catalyst bed

Micro thrusters, also referred to as micro-newton thrusters, are thrusters having a wide range of force capabilities in the micro-newton range. Micro thrusters are used in CubeSats, tiny robots and drones, micro- and nanosatellites, spacecraft attitude control, etc. The aim of the present study is to develop a monopropellant thruster (MPT) for 90 % concentrated hydrogen peroxide (H2O2). A rotatory evaporator was used to obtain a 90 % H2O2 form the available concentration of 50 % H2O2. The next stage involved the designing of the monopropellant thruster utilizing the NASA CEA rocket performance algorithm and the derived mathematical model. The test facility was constructed which includes a thruster, a feed line system, a static test firing platform, a data collecting system and the measuring sensors. To measure the thrust of the thruster, a pendulum thrust mechanism was constructed and fabricated into an experimental test stand. Lastly, the MPT chamber, which uses stainless steel mesh, was prepared and packed with the biomass-based catalyst. The 14 bar injection pressure conditions were used for the fire tests. With a mass flow rate of 5 g/s and a decomposition pressure of 10 bar, the theoretical specific impulse (Isp) was calculated to be 1993.2 Ns/kg at an O/F ratio of 0.002. The test firing with 90 % concentrated hydrogen peroxide gave an experimental specific impulse of 794.5 Ns/kg, which gives efficiency of 39.9 %.

Influence of Aluminum Content and Agglomerates Initial Velocity on Erosion in Solid Rocket Motor

Abstract Despite the aluminized propellants offering a high specific impulse, the challenge of nozzle erosion adversely impacts the rocket's performance and its potential for reusability. This study presents a numerical model aiming to predict the mechanical erosion of the propulsion chamber nozzle. The model employs an Eulerian/Lagrangian approach to simulate the complexity of the flow field within the rocket combustion chamber and the interactions between the continuous phase and particles. The model also includes a simplified representation of the aluminum particle combustion process, besides the consideration of secondary breakup phenomena in liquid droplets. Experimental and numerical data from the literature were used to validate the numerical model. Subsequently, the model was utilized to explore the impacts of increasing propellant aluminum content and varying particles' injection velocities on the nozzle mechanical erosion. The outcomes indicated that higher aluminum content leads to a 4-10% increase in nozzle erosion compared to the 15% content case. Furthermore, the aluminum particles tend not to fully burn within the combustion chamber and contribute to nozzle erosion. Lastly, particles with higher initial velocity at the inlet of the combustion chamber increase the nozzle mechanical erosion despite the observed decrease in incident mass flux.

Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets.

This study compares propellant fuels' specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with Ammonium Perchlorate (NH 4ClO 4) as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Test results with the composition of CBK2 (Coconut Shell + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket's thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

Design optimization of a low-cost three-stage launch vehicle with modular hybrid rocket motors

This article investigates the impact of modular propulsion system design on the performance and cost of a three-stage hybrid rocket. Furthermore, it conducts a multi-objective optimization of unit payload cost, take-off mass, and payload mass ratio, considering factors such as the number of motors and layout considerations. The optimization design scheme for the three-stage hybrid rocket is divided into four cases. In the first case, each stage is equipped with a fixed single motor, and each stage is independently optimized without modular design. The second case considers the use of multiple motors in the first and second stages, still without modular design. The third case also involves multiple motors in the first and second stages, but all motors in each stage have identical parameters except for the nozzle expansion ratio, implementing a modular design. In the fourth case, the number and layout of the motor design method are the same as those in the third case, with independent optimization in the third stage using partial modular design. The results indicate that the unit payload cost of the multi-motor non-modular design case can be reduced by 13.12% compared to the single-motor non-modular design case. Within the modular case, the full modular design case is slightly inferior to the partial modular design case. Based on the above data, it can be concluded that the first and second stages of modular rockets offer the best performance and the lowest cost.

Optimizing and designing upper-stage separation spring of lifting-body rocket

Because the upper-stage separation of a lifting-body rocket is difficult under the low-altitude and high-speed flight environments, an upper-stage separation scheme based on the separation spring is proposed. The integration of ADAMS and MATLAB realizes the multi-body dynamics simulation analysis of upper-stage separation. The experimental design method is used to carry out the target simulation and calculation of the upper-stage separation process, with the separation condition deviation and the separation spring deviation considered at the same time, thus ensuring the reliability of the separation scheme. In order to further optimize the separation scheme, an upper-stage separation optimization model is established with the Isight software, and the mass of separation spring is optimized with the multi-island genetic algorithm and significantly reduced by 48.87% under the condition that the upper-stage separation process satisfies all constraints, thereby effectively improving the performance of the lifting-body rocket.

Optimization of Nozzle Design for Weight Reduction using Variable Wall Thickness

In various fields, including rockets, turbines, and engines, the critical role played by nozzles in achieving optimal performance is underscored. Within the domain of rocketry, particular attention is given to the weight of the nozzle, where even the slightest modifications in weight can lead to a profound impact on rocket performance. In this research, an endeavor is made to comprehensively investigate the intricate interplay of nozzle pressure ratio, nozzle thickness, and Mach number, to unveil the effects of these parameters on nozzle deformation, the development of equivalent stress, and the pivotal factor of safety. Material optimization strategies for a wide array of flow and geometrical parameters are also explored within the framework of this study. With each parameter considered at four levels, an exploration is conducted across Mach numbers of 1.5, 2.0, 2.5, and 3.0, alongside Nozzle Pressure Ratios of 2, 4, 6, and 8. The meticulous analysis of nozzle thicknesses at levels of 1 mm, 2 mm, 3 mm, and 4 mm is carried out. The computational fluid dynamics (CFD) analysis is systematically executed across all cases, with resultant pressure profiles on the internal surfaces of the nozzle serving as essential boundary conditions for static structural analysis. The presentation of findings entails a comprehensive discussion of the factor of safety values for all parameter combinations. Significantly, it is observed from our results that a direct correlation exists between nozzle thickness and factor of safety, with an increase in nozzle thickness corresponding to an enhancement in factor of safety.

Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets

Background: This study compares propellant fuels’ specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods: The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with NH4ClO4 as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Results: Test results with the composition of CBK2 (Coconut Shell + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. Conclusions: The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket’s thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

Study on the dynamic numerical simulation of flow and combustion in hybrid rocket motors based on a discrete phase model

Numerical simulation based on the discrete phase model and dynamic mesh model can more accurately simulate the working process for hybrid rockets, including the motion and evaporation process of particles and regression process of the fuel grain. The dynamic numerical simulation of combustion and flow in hybrid rocket motors based on a discrete phase model is carried out. Aluminum and aluminum hydride are added to the internal flow field as discrete droplets. Then, a hybrid rocket is designed with 95 % hydrogen peroxide as oxidizer and 38 % aluminum hydride+20 % aluminum+42 % hydroxyl-terminated polybutadiene as the fuel and used to verify the model by firing test. A high degree of agreement between experiment and simulation findings is found. The average deviations for thrust and pressure are 3.6 % and 5.9 %, respectively. It offers the required backing for using the simulation method to investigate the flow and combustion process for the hybrid rockets. The results also demonstrate the transient characteristics of the inner flow field, including the separation and merging of vortexes in the pre-chamber. It is found that a peak of the regression rate appears at the front of the grain and moves backward. This is mainly because the vortex appears here and enhances its heat transfer. Through the discrete phase model, it is worth noting that the aluminum droplet content gradually decreases at the nozzle over time, which is consistent with the firing test. Further research on the impact of aluminum hydride content on the performance of the hybrid rockets has been carried out. The results show that the higher aluminum hydride content can lead to better performance for the hybrid rockets. The thrust, pressure, specific impulse, and characteristic velocity enhance with the increase of aluminum hydride content. An interesting discovery is that the average regression rate is maximum at 48 % aluminum hydride content. This research innovatively couples dynamic mesh and discrete phase models, which can help better understand the working process for hybrid rockets and provide important guidance for the selection of fuel formula.

Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets

Background: This study compares propellant fuels' specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods: The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with NH4ClO4 as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Results: Test results with the composition of CBK2 (Coconut Shell + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. Conclusions: The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket's thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

Predicting Performance and Roll-Rate Minimization of a High-Powered Model Rocket

The Collegiate Rocket Launch is a competition sponsored by the Wisconsin Space Grant Consortium which challenges student teams from colleges throughout the state of Wisconsin to design, build, and launch a high-powered model rocket. Teams are scored based on quality of technical reports, project management, outreach, and the competition flight. The 2022 competition challenged teams to create an active control system that minimized the roll rate of the rocket while in flight, as well as reach but not exceed a target altitude of 3,500 feet. Marquette Rocketry is proud to have placed first in the competition and given the title mission grand champion.

Photosynthesis, Yield and Quality in Wild Rocket (Diplotaxis tenuifolia L.) under Photoluminescent Greenhouse Covers

Manipulation of light spectral composition is a useful tool to drive morphological, physiological and metabolic responses in several crops, ultimately improving yield and quality. Novel materials for greenhouse covering are being developed in order to make a better use of the available sunlight: among these are the cover films or panels incorporating fluorescent additives which are able to convert UV solar radiation into visible light. In this research, we compared the physiological traits and the agronomical performance of wild rocket grown in pots in the winter–spring season, under four different greenhouse prototypes covered with poly-methyl-methacrylate (PMMA)-based panels. PMMA panels doped at 3% (Dop3) or 7% (Dop7) w/w with a blend of rare-earth elements (partially converting the solar UV radiation to red and blue wavelengths) were compared with an undoped (UD) and a whitewashed (WW) PMMA greenhouse. The rocket yield was higher in Dop3 (+30%), while it was unaffected in Dop7 and lower in WW (−39%), compared to the control (6.06 kg m−2). The leaf greenness decreased while both the ABTS and the hydrophilic antioxidant activities increased under the doped and the whitewashed greenhouses. The Dop3 treatment provided the best results in terms of yield and quality of greenhouse wild rocket in winter–spring cycle. However, the analysis of OJIP kinetics of chlorophyll fluorescence revealed that the main factor affecting the photosynthetic performance was the light intensity inside each greenhouse rather than the modulation of light spectrum, because of the different shading properties of the doping and whitewashing treatments. Although these results did not allow us to distinguish between the combined effects of shading and light spectrum modulation, the use of photoluminescent covers can be foreseen as a promising innovation in greenhouse horticulture.

Investigation of Flight Performance of Notched Delta Wing Rockets on Different Types of Nose Cones

In this study, four solid fuel model rockets with conical, parabolic, power and haack series nose cones that can carry 4 kg payload at an average altitude of 3 km were designed in the OpenRocket program. Later, the notched delta fin model was mounted on these designed model rockets. The effects of this fin model on the changes in the speed, stability, acceleration, weight and altitude of the rockets were analyzed numerically in the OpenRocket program. As a result of the analysis, it was determined that the conical nose rocket showed the worst flight performance and the Haack series nose cone rocket model showed the best performance. When used with the notched delta fin of the Haack series model, it was determined that the rocket's altitude increased by 7.67%, and its speed increased by 1.83%, but decreased by 1.2% in mach number, 0.6% in weight, 0.3% in acceleration, and 4.5% in stability. As a result, it was seen that it would be beneficial to consider the nose cone and fin together when evaluating the flight performance of the rocket. The results obtained in the study have shown that the notched delta fin model can be used experimentally in defense industry and model rocket applications and the studies can be advanced.

Experimental validation of rotating detonation for rocket propulsion

Space travel requires high-powered, efficient rocket propulsion systems for controllable launch vehicles and safe planetary entry. Interplanetary travel will rely on energy-dense propellants to produce thrust via combustion as the heat generation process to convert chemical to thermal energy. In propulsion devices, combustion can occur through deflagration or detonation, each having vastly different characteristics. Deflagration is subsonic burning at effectively constant pressure and is the main means of thermal energy generation in modern rockets. Alternatively, detonation is a supersonic combustion-driven shock offering several advantages. Detonations entail compact heat release zones at elevated local pressure and temperature. Specifically, rotating detonation rocket engines (RDREs) use detonation as the primary means of energy conversion, producing more useful available work compared to equivalent deflagration-based devices; detonation-based combustion is poised to radically improve rocket performance compared to today’s constant pressure engines, producing up to 10% increased thrust. This new propulsion cycle will also reduce thruster size and/or weight, lower injection pressures, and are less susceptible to engine-damaging acoustic instabilities. Here we present a collective effort to benchmark performance and standardize operability of rotating detonation rocket engines to develop the RDRE technology readiness level towards a flight demonstration. Key detonation physics unique to RDREs, driving consistency and control of chamber dynamics across the engine operating envelope, are identified and addressed to drive down the variability and stochasticity observed in previous studies. This effort demonstrates an RDRE operating consistently across multiple facilities, validating this technology’s performance as the foundation of RDRE architecture for future aerospace applications.

Experimental and numerical investigation on design parameters of slot nozzles

Slot nozzles are altitude compensation nozzles. In the overexpansion mode, slots in the expansion section can affect the flow separation and improve the thrust. In this study, cold-flow experiments and numerical simulation are performed to study the design parameters of slot nozzles. The effects of slot location, number, and vented area on the flow field, thrust, and average specific impulse of the first-stage rocket are investigated. The results show that the thrust increases due to two reasons: the advance of the separation position and the rise of wall pressure in the recirculation zone. Moreover, the slot location is more important than other parameters in the design of slot nozzles. Only when a slot is close to the flow separation point, the influence of the slot on the flow field is apparent, thus, the thrust significantly and efficiently increases. The performance of the first-stage rocket can be significantly improved by replacing the traditional bell nozzles with slot nozzles. The average specific impulse increase of the space shuttle solid rocket booster and first-stage rocket of the nano launch vehicle (NLV) is 5.1 s and 7.9 s, respectively; further analysis of NLV shows that its payload can be increased by 16.1%.

THE EFFECT OF FLIGHT PARAMETERS UNCERTAINTY ON THE R-HAN 300 ROCKET FALL POINT

The impact point is the main parameter of the performance of an artillery rocket. During its flight, a rocket experiences uncertainty of some parameter’s value, which is caused by the condition of the rocket when it exits the launcher, the ever-changing atmospheric conditions, and fabrication imperfections. This disturbance will cause deviation on the impact point. This study aims to determine the deviation of the impact point of an artillery rocket due to the value uncertainty of several parameters that affect its flight trajectory, and this analysis will be used as a reference for designers and users in further design processes. Six degrees of freedom simulation calculates the rocket's impact point and deviation. The results showed that to minimize the impact point’s deviation, the quality of the launcher fabrication must be improved, predicting aerodynamics with CFD and firing the rocket at low wind disturbances.

On the effect of pitch and yaw angles in oblique impacts of small-caliber projectiles

A terminal ballistic analysis of the effects of 7.62 mm × 51 AP P80 rounds on inclined high-strength armor steel plates is the focus of the presented study. The findings of an instrumented ballistic testing combined with 3D advanced numerical simulations performed using the IMPETUS Afea® software yielded the conclusions. The experimental verification proved that slight differences in the pitch-and-yaw angles of a projectile upon an impact caused different damage types to the projectile’s core. The residual velocities predicted numerically were close to the experimental values and the calculated core deviations were in satisfactory agreement with the experimental results. An extended matrix of the core deviation angles with combinations of pitch-and-yaw upon impact angles was subsequently built on the basis of the numerical study. The presented experimental and numerical investigation examined thoroughly the influence of the initial pitch and yaw angles on the after-perforation projectile’s performance.

Effects of weak connecting structure on penetration performance of free-flight rocket

The strength of connection between the rocket engine and the warhead has a certain relationship with the warhead’s ballistic characteristics and structural response. In this paper, the effects of weak connecting structure on the penetration performance of the free-flight (unguided) rockets is studied by numerical simulation and experiment. The oblique penetration of free-flight rockets with different connecting structures into armour plates is numerically simulated, the rockets’ deflection in the penetration process is analysed, and the effects of the weak connecting structure on the warhead’s structural integrity are evaluated according to the deformation of the warhead tail. The oblique penetration test of free-flight rocket is designed, and the simulation results under low impact velocities are verified. The research shows that the influence of the length of the weak connecting structure on the tail deformation of the warhead varies at different impact velocities. It is better to use direct connection between the engine and the warhead at high impact velocities, and the weak connecting structure performs better at low impact velocities.

PERFORMANCE OF SMALL SIZE AEROSPIKE. ENGINE

In this paper we are showing how aerospike nozzle can play an important role in rocket INDUSTRY as compare to other nozzles. We are using New rocket designs and increasing the performance of the current rockets But, the aerospike nozzle concept that has been under development since the 1950s is yet to be utilized on a launching. Due to its ability to adjust the environment by altering the outer jet boundary, the aerospike nozzle has better performance compared to present day others nozzle. An aerospike nozzle is designed for 20 bar pressure ratio to improve the performance of the aerospike nozzle for various conditions, some important design parameters and their performances were studied for low temperature flow conditions.