The rapid developments in micro-technologies and the introduction of modularity and standardization into system designs, present significant opportunities for cost reduction in the design and development of satellite systems. However, the high cost of space launch has become a major hindrance to capitalizing on these opportunities. Therefore, seeking appropriate launch opportunities and reducing launch costs might contribute to further growth of the space market. This paper focuses on the analysis of dedicated launch costs factoring in the effect of launch reliability, which in return, can enable the optimization of system designs. Applying a value-centric architecture, system characteristic space is introduced as the design space to define the characteristics of different systems. Based on our launch vehicle database, the launch cost and reliability of different families of launch vehicles are investigated, where the reliability is calculated using a modified two-level Bayesian analysis. The factors of launch cost and reliability are subsequently integrated into the expected launch cost, acting as the objective function for the analysis and optimization process associated with the manufacturing cost of satellites. Through reviewing and redesigning a few classical launch cases, the effectiveness and applicability of the design architecture proposed are validated.
Liquid-propellant pulsed plasma thrusters (LP-PPTs) perform better than conventional solid-propellant (typically PTFE) PPTs because the use of an LP can eliminate the problems of late-time ablation and particulate emission associated with solid propellants. In the present study, the performance and characteristics of a prototype LP-PPT are investigated using a range of nozzle and electrode configurations. Two types of conical nozzles were tested: a ceramic nozzle with an annular anode placed at the nozzle tip and a monolithic-anode nozzle made of stainless steel. The area ratio ε, divergent angle θ, and cavity length L of the ceramic nozzle were varied from 10 to 30, 10° to 40°, and 5 to 20 mm, respectively. Ethanol was used as the propellant. The thrust measurements showed that the LP-PPT prototype demonstrated superior performance when it was fitted with the embedded annular anode ceramic nozzle than when it was fitted with the monolithic-anode nozzle. Among all of the tested nozzle configurations, the highest performance was observed for a ceramic nozzle with ε = 30, θ = 20°, and L = 10 mm, which yielded an impulse bit of 167 µNs, a specific impulse of 1,150 s, a thrust efficiency of 5.9%, and a thrust–power ratio of 11 µNs/J at a capacitor-stored energy of 14 J and a propellant mass shot of 14.8 µg.
The flight dynamics subsystem (FDS) of a geostationary satellite ground control system conducts estimates of the orbital states, keeping within its mission box, and calculating the satellite operational parameters to be uploaded. A new FDS has been designed and implemented by dividing it into core- and bus-dependent modules for a recent Korean geostationary satellite ground control system. The core modules include orbit determination and prediction, event prediction, and station-keeping and relocation (SKR) planning; whereas the bus-dependent modules include fuel accounting, thruster modeling and maneuver reconstructions of SKR, and calculate all of the bus-dependent parameters. In particular, separate designs based on the components of the SKR planning and maneuver reconstruction using thruster modeling allow the system to be reusable and replaceable. We also designed and implemented a conjunction analysis tool and collocation control and monitoring units for multiple-satellite control. A FDS database has been developed and is managed using SQLite, which is freely distributed. The FDS is easy to develop and operate thanks to a novel separation concept introduced for the core-platform and spacecraft bus-dependent modules. It is the same concept used for operation and the graphical user interfaces (GUIs) applicable to the FDS. All of the FDS modules for new geostationary satellites have been validated through function and performance testing, and have proven to work successfully after the launch of satellites.
The effects of a recess on co-flowing planar jets under supercritical pressure are numerically studied. Two-dimensional hybrid LES/RANS simulations are performed in a wide range of recess lengths and injection momentum flux ratios, which are important design parameters for liquid rocket engine injectors. The present study showed that confinement effects of the near-injector flowfield by applying a recess, suppress outer jet spreading and thus enhances the penetration of the outer jet flow into the inner jet. The enhanced penetration of the outer jet flow results in the appearance of flapping motions in the inner jet. Low-frequency oscillations corresponding to the flapping motions clearly appear in the case of recessed injectors. Moreover, the confinement effects promote interactions between vortex structures resulting in vortex breakdown. Consequently, the inner-jet length is shortened, indicating an improvement in mixing when a recess is applied. The present results also show that the inner-jet length deceases as the recess length increases, and the effects of a recess remarkably appear at higher momentum flux ratios. This is explained by the vortices generated behind the post lip that is strengthened as the result of increased velocity ratio.
In bi-propellant thrusters, impinging type injectors are widely used to deliver propellants to a combustion chamber. By impinging the jet streams of fuel and oxidizer, the spray spreads while the two liquids mix. To design the injectors, several correlations related to injection conditions have been proposed (e.g., Rupe factor), and practically utilized over the last half-century. However, the physical meanings of the past correlations are not well understood, because the essential scale of the spray structure is elusive. In this paper, we derive the global length scale of the spray produced by impinging injectors of unlike doublet, fuel-oxidizer-fuel triplet, and oxidizer-fuel-oxidizer triplet in a consistent manner. The unified length scale is found representing the spray width ratio of oxidizer to fuel evidenced by comprehensive cold-flow tests including several past studies, covering various parameters such as injector types, nozzle diameters, physical properties of working liquids, and injection velocities. Finally, we clearly provide the physical meaning based on practical correlations in a phenomenological sense.
This study is focused on the effect of boattail angle on the pressure distribution and drag force of axisymmetric afterbodies under low-speed conditions. Experiments were conducted on three conical boattails with angles of 10°, 14° and 20°. The diameter-based Reynolds number is approximately 4.3 × 104 under the experimental conditions. Two types of flows, a fully-attached flow (β = 10°) and a flow with a separation bubble (β = 14°, 20°), were observed. The aerodynamic drag measurements were conducted using both a strut-supported model and a support-free (levitated) model. The results show that boattail model with the angle of 20° has a relatively large effect on the pressure distribution. The pressure drag resulting from pressure distribution on the vertical plane indicates that the model with a boattail angle of 14° has the lowest drag. A good trend in agreement between afterbody drag (measured using pressure taps) and total aerodynamic drag (measured using the levitated system) was obtained. The effect of strut support on pressure distribution at different polar angles is also explained in this study.
In this study, a linear reduced-order model of flow fields around a NACA0015 is constructed based on time-resolved particle image velocimetry (PIV) data. The PIV data were obtained at the chord Reynolds number of 6.4 × 104, and angles of attack from 11° to 20°. Proper orthogonal decomposition (POD) analysis is employed for the PIV data and the degrees of freedom are reduced by truncating the POD modes. Next, a linear model of the POD modes is constructed using the least-squares method based on the POD-mode time histories. Although the estimated POD modes initially reproduce original data, they gradually attenuate and converge to zero. This behavior is also supported by the eigenvalue analysis results of the model's coefficient matrices. In addition, the behavior of the low-order (more energetic) POD modes was reproduced better than that of the high-order (less energetic) POD modes. These results imply that the temporal fluctuation of large vortex structures has strong linearity and is not significantly affected by noise included in data. The former insight is also supported by the fact that the behaviors of POD modes were reproduced well in the case of high angle of attack.
An adaptive sliding mode control to stabilize the attitude of a bias momentum satellite with a time delay via two wheels is proposed. Stabilizing the attitude of a rigid body via two control torques is an under-actuated control problem, and belongs to the class of systems that are controllable but cannot be asymptotically stabilized via continuous state feedback because Brockett's necessary condition is not satisfied. The adaptive method combined with the sliding mode control estimates the time delay contained in the system by sensing the difference between the attitude predicted and the measured one at each sampling time and compensates for the time delay by predicting the current state using the past measured attitude and angular velocity. Provided that external disturbances and modeling uncertainties in the satellite moments of inertia are absent, the validity of the proposed adaptive time delay estimated sliding mode control for attitude control of a bias momentum satellite is verified through numerical simulations.