• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.48 no.11
• /
• pp.873-880
• /
• 2020

In this paper, we propose a guidance filter to estimate line of sight rate with strapdown seeker measurements and INS(Inertial Navigation System) information. The measurements of proposed guidance filter consisted of the LOS(Line of Sight) and relative position that can be calculated with the seeker's measurements, INS information and known target position, also the filter is based on an asynchronous filter to use outputs of the two sensors that are out of synchronous and period. Through the proposed filter, we can reduce the effect on parasitic loop that can be caused by using large time delay seeker and improve the estimation performance.

• Journal of the Korean Society for Precision Engineering
• /
• v.15 no.6
• /
• pp.97-106
• /
• 1998

In this Paper, a fuel minimizing closed loop explicit inertial guidance algorithm for orbit injection of a rocket is developed. In the formulation, the fuel burning rate and magnitude of thrust are assumed constant. The motion of rocket is assumed to be subject to the average inverse-square gravity, but negligible effects from atmosphere. The optimum thrust angle to obtain a given velocity vector in the shortest time with minimizing fuel consumption is first determined, and then the additive thrust angle for targeting the final position vector is determined by using Pontryagin's maximum principle. To establish real time processing, many algorithms of onboard guidance software are simplified. The explicit guidance algorithm is simulated on the 2nd-stage flight of the N-1 rocket developed in Japan. The results show that the explicit guidance algorithm works well in the presence of the maximum $\pm$10% initial velocity and altitude errors, and exhibits better performance than the open-loop program guidance. The effects of the guidance cycle time are also examined.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.40 no.10
• /
• pp.853-861
• /
• 2012

This paper considers one of the explicit guidance algorithms, which has been proposed by Jaggers, to determine the closed-loop guidance algorithm for upper stages of a 3-staged space launch vehicle. Its commanded thrust vector is closer to the optimal solution when compared with that obtained by using the well-known Powered Explicit Guidance (PEG), which has been developed through the Space Shuttle program. Its performance is evaluated here by applying for guidance of the launcher during the second and third stages. Furthermore, to generate more precise guidance commands, it is attempted not to use the approximate formulas for the derivation of the original guidance law, and it is shown that performance is improved in comparison with the original.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.39 no.6
• /
• pp.505-516
• /
• 2011

The concept of the IGC(Integrated Guidance and Control) has been introduced to overcome the performance limit of the SGC(Separated Guidance and Control) loop. A new type of IGC with impact angle constraint has been proposed in this paper. Angle of attack, pitch angle rate, pitch angle and line of sight angle are considered as state variables. A controllability analysis and equilibrium point analysis have been carried out to investigate the control characteristic of the prposed IGC. The LQR(Linear Quadratic Regulator) has been adopted for the control law and detailed explanations about the adoption has been provided. The performance comparison between the IGC and the SGC has been carried out. The result of numerical simulations shows that the IGC guarantees better guidance performance than the SGC when the agile maneuver is needed for a specific guidance geometry.

• International Journal of Aeronautical and Space Sciences
• /
• v.17 no.3
• /
• pp.366-377
• /
• 2016

This paper presents a vision-based guidance method that allows a fixed-wing aircraft to orbit around a target at a given radius. The guidance method uses a simple formula that regulates a relative side-bearing angle estimated by a vision system. The global asymptotic stability of the associated guidance law is demonstrated, and a linear analysis is performed to facilitate the proper selection of the relevant control parameters. A flight experiment is presented to demonstrate the feasibility and performance of the proposed guidance method.

• Journal of the Korean Society for Industrial and Applied Mathematics
• /
• v.19 no.3
• /
• pp.327-364
• /
• 2015

In this paper, a generalized guidance law with an arbitrary pair of guidance coefficients for impact angle control is proposed. Under the assumptions of a stationary target and a lag-free missile with constant speed, necessary conditions for the guidance coefficients to satisfy the required terminal constraints are obtained by deriving an explicit closed-form solution. Moreover, optimality of the generalized impact-angle control guidance law is discussed. By solving an inverse optimal control problem for the guidance law, it is found that the generalized guidance law can minimize a certain quadratic performance index. Finally, analytic solutions of the generalized guidance law for a first-order lag system are investigated. By solving a third-order linear time-varying ordinary differential equation, the blowing-up phenomenon of the guidance loop as the missile approaches the target is mathematically proved. Moreover, it is found that terminal misses due to the system lag are expressed in terms of the guidance coefficients, homing geometry, and the ratio of time-to-go to system time constant.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.36 no.8
• /
• pp.747-752
• /
• 2008

We propose a guidance law based on pursuit-evasion game solutions, and analyze its performance. The game solutions are obtained from the pursuit-evasion game solver developed by Tahk. The initial value of the game solution is used for guidance, and then the pursuit-evasion game is solved again at the next time step. In this respect, the proposed guidance laws are similar to the approach of model predictive control. The proposed guidance method is compared with proportional navigation guidance for a pursuit-evasion scenario, in which the evader always tries to maximize the capture time. According to the comparison, it has larger a capture set than ones of proportional navigation guidance law.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.47 no.8
• /
• pp.574-581
• /
• 2019

This paper investigates the effects of thrust misalignments and offsets of the lateral center of gravity of a space launch vehicle on its guidance performance. Sensitivity analysis and Monte Carlo simulations are applied to analyze their effects by computing changes in orbit injection errors when including the error sources. To compensate their effects, the attitude controller including an integrator additionally and the Steering Misalignment Correction (SMC) routine of the Saturn V are considered, and then Monte Carlo simulations are performed to evaluate their performances.

• International Journal of Aeronautical and Space Sciences
• /
• v.11 no.2
• /
• pp.110-117
• /
• 2010

We propose guidance laws based on a pursuit-evasion game. The game solutions are obtained from a pursuit-evasion game solver developed by the authors. We introduce a direct method to solve planar pursuit-evasion games with control variable constraints in which the game solution is sought by iteration of the update and correction steps. The initial value of the game solution is used for guidance of the evader and the pursuer, and then the pursuit-evasion game is solved again at the next time step. In this respect, the proposed guidance laws are similar to the approach of model predictive control. The proposed guidance method is compared to proportional navigation guidance for a pursuit-evasion scenario in which the evader always tries to maximize the capture time. The capture sets of the two guidance methods are demonstrated.

• Aerospace Engineering and Technology
• /
• v.11 no.1
• /
• pp.169-177
• /
• 2012

This paper considers the explicit guidance algorithm to determine the closed-loop guidance law applicable to upper stages of a given space launch vehicle. It has the advantage of very simple forms derived from the flat earth assumption, which is appropriate for its on-board application. However the simple time-to-go prediction equation produces the degraded guidance performance of the launcher because of its inaccuracy. To overcome the problem, the elaborate prediction equations, which have been employed in Saturn and H-II, are attempted here. Finally, the simulation results show that the simple guidance approach requires the more accurate time-to-go prediction and gravity integrals for its broad application.