Vehicle Ascent Research Articles

Launch Vehicle Ascent Computational Fluid Dynamics for the Space Launch System

This paper will discuss the use of computational fluid dynamics (CFD) for the Space Launch System (SLS) program to model the ascent phase of flight. The ascent phase begins shortly after the vehicle clears the launch tower and extends to the first staging event. To model SLS’s ascent, over 1000 numerical solutions of the Navier–Stokes equations were solved, and this analysis has been repeated for five different SLS configurations. To manage this demanding ascent CFD task, the SLS program has developed the Computational Aerosciences Productivity & Execution software. The paper also discusses some of the ways that CFD and high-end computing have advanced in the last decade and offers some comparisons to CFD used in the Space Shuttle Program.

Investigating the rocket base flow characteristics and thermal environment of different nozzle configurations considering afterburning

During launch vehicle ascent, the base flow regimes of multi-nozzle rockets become increasingly complex owing to plume interactions. We developed a numerical method using the hybrid RANS/LES approach and the discrete ordinate method to accurately estimate the base thermal environment. Moreover, the finite-rate chemical kinetics are employed to calculate afterburning reactions. The validity of the numerical method is established by achieving a good agreement between the numerical results and wind tunnel data. Subsequently, the reacting and non-reacting flows of the three types of nozzle configurations were simulated at six different altitudes. The numerical results reveal that the plume interactions become stronger with the increase in flight height and total number of nozzles, and the maximum Cp of the four-nozzle rocket was 34.75% and 56.52% higher than that for three- and two-nozzle rocket at 40 km, respectively. After considering the afterburning effect, the maximum temperatures of the two-, three-, and four-nozzle rockets increased by 10.16%, 7.69%, and 5.43%, respectively, owing to the difference in plume collisions. With increasing altitude, the rocket base heating rate distribution changed significantly. The peak heat flux of the rocket base initially increases before decreasing with the maximum value of 147.32 kW/m2, 872.75 kW/m2, and 1395.80 kW/m2 for two-, three-, and four-nozzle rockets, respectively.

A novel linear uncertainty propagation method for nonlinear dynamics with interval process

Interval process is a preferable model for time-varying uncertainty propagation of dynamic systems when only the range of uncertainties can be obtained. However, for nonlinear systems, except for Monte Carlo (MC) simulation, there are still few efficient uncertainty propagation methods under the interval process model. This paper develops a non-intrusive and semi-analytical uncertainty propagation method, named the “convex model linearization method (CMLM),” by constructing a linearization formulation of a nonlinear system in a non-probabilistic sense. First, the criterion to evaluate the difference between the original system and the linearization formulation is derived, represented by the discrepancy of middle point, radius and correlations of response. By minimizing these three parameters, the coefficients of linear equations will be optimized to obtain the linearization formulation of the original system. Then, analytical equations are built to calculate uncertainty response under the interval process, without time-consuming analysis of the original system. To further improve the efficiency of the linearization process, Chebyshev polynomial is introduced to approximate the nonlinear dynamic analysis. Two numerical examples of duffing oscillators and vehicle rides are set to test the proposed CMLM. Compared to the MC method, with comparable uncertainty response precision, the CMLM just needs 1–10% times of dynamic analyses of the nonlinear system. Furthermore, a practical launch vehicle ascent trajectory problem with black-box dynamics is solved by, respectively, the CMLM and MC method. The results verify the capacity of the CMLM to deal with black-box problems and show that the CMLM performs better in terms of accuracy, efficiency and robustness.

Convex Optimization of Launch Vehicle Ascent Trajectory with Heat-Flux and Splash-Down Constraints

This paper presents a convex programming approach to the optimization of a multistage launch vehicle ascent trajectory, from the liftoff to the payload injection into the target orbit, taking into account multiple nonconvex constraints, such as the maximum heat flux after fairing jettisoning and the splash-down of the burned-out stages. Lossless and successive convexification methods are employed to convert the problem into a sequence of convex subproblems. Virtual controls and buffer zones are included to ensure the recursive feasibility of the process, and a state-of-the-art method for updating the reference solution is implemented to filter out undesired phenomena that may hinder convergence. A pseudospectral discretization scheme is used to accurately capture the complex ascent and return dynamics with a limited computational effort. The convergence properties, computational efficiency, and robustness of the algorithm are discussed on the basis of numerical results. The ascent of a VEGA-like launch vehicle toward a polar orbit is used as a case study to discuss the interaction between the heat flux and splash-down constraints. Finally, a sensitivity analysis of the launch vehicle carrying capacity to different splash-down locations is presented.

Convex Approach to Three-Dimensional Launch Vehicle Ascent Trajectory Optimization

This paper deals with the optimization of the ascent trajectory of a multistage launch vehicle, from liftoff to the payload injection into the target orbit, considering inverse-square gravity acceleration and aerodynamic forces. A combination of lossless and successive convexification techniques is adopted to generate a sequence of convex problems that rapidly converges to the original problem solution. An automatic initialization strategy is proposed to make the solution process completely autonomous. In particular, a novel three-step continuation procedure is developed and proved to be more efficient than simpler strategies. This approach relies on the solution of intermediate problems, which either neglect atmospheric drag or fix the time-lengths of the launch vehicle ascent phases, that are solved in succession, gradually passing from easier instances of the optimization problem to the originally intended problem. State-of-the-art techniques to deal with such a complex problem are adopted to enhance the convergence rate, including safeguarding modifications, such as virtual controls and an adaptive trust region. To assess the validity of the proposed approach in a practical scenario, numerical results are presented for two representative practical applications, using as reference a Falcon 9 launch vehicle.

ISRU technology deployment at a lunar outpost in 2040: A Delphi survey

The purpose of this study was to deploy a Delphi expert elicitation methodology to better understand the technical and policy challenges facing the development of a sustainable lunar outpost in 2040, including the types and scale of In-Situ Resource Utilisation (ISRU) deployment. We used a three-round Delphi survey with an open first round and specific questions in later rounds using a four-point Likert scale and two ranking exercises to assess energy technologies and inhibiting factors. In order to provide more certainty to our potential participants regarding their input, and boost engagement, the study deployed a three-round approach that was communicated to our potential participants and decided ex-ante. Potential participants were identified from the literature and academic networks as those who had made significant contributions to the fields of: ISRU technologies, space architecture, space-qualified power systems, and space exploration. The study identified around 20 major themes of interest for researchers in the first round and asked participants to rate their agreement with a number of statements about a hypothetical lunar outpost in 2040. From the group responses, we identified three major technical challenges for the development of a lunar outpost in 2040; developing high power energy infrastructure, lander and vehicle ascent capacity, and mission architectures and technical approaches. We also identified three major policy challenges for the development of a lunar outpost in 2040: (i) US and global political instability, (ii) possibility of an extended timeframe for the first lunar landing, and (iii) political distaste for nuclear energy in space. The group was uncertain about the precise energy mix at the outpost as a result of uncertainty regarding electrical loads, but there was general agreement that solar PV would be a significant contributor. Whether nuclear power sources might play a useful role proved to be very uncertain, with some participants noting a political distaste for space nuclear power systems. However, the proposition gained two votes in each ranking position, suggesting it has a flat distribution including both supporters and detractors.

Ascent Trajectory Optimization for Hypersonic Vehicle Based on Improved Chicken Swarm Optimization

Trajectory optimization problem for hypersonic vehicles has received wide attention as its high speed and large flight range. The strong nonlinear characteristic of the ascent phase aerodynamics makes the trajectory optimization problem difficult to be solved by the optimal control theory. In this paper, an improved chicken swarm optimization (ICSO) algorithm is proposed to optimize the hypersonic vehicle ascent trajectory. To overcome the obstacle of premature convergence, three improvement strategies are put forward. To be specific, the updating laws of roosters are modified by the average position of roosters, and the difference of the optimal solution between two adjacent iterations is used to calculate the mutated particle instead of the gradient. Meanwhile, the uniform mutation operator is used to get rid of the local minimum. The convergence analysis of the proposed ICSO is provided subsequently. To handle constraints, an improved adaptive penalty method is put forward. The comparison results show that the proposed ICSO outperforms CSO and PSO on benchmark functions of CEC2014. Finally, the trajectory optimization results for a generic hypersonic vehicle, in compare with the open-source optimization software PSOPT, are put forward to demonstrate the feasibility and effectiveness of the proposed method. The results of 50 independent runs show that the improved adaptive penalty function method is effective in constraints handling.

A Multiconstrained Ascent Guidance Method for Solid Rocket-Powered Launch Vehicles

This study proposes a multiconstrained ascent guidance method for a solid rocket-powered launch vehicle, which uses a hypersonic glide vehicle (HGV) as payload and shuts off by fuel exhaustion. First, pseudospectral method is used to analyze the two-stage launch vehicle ascent trajectory with different rocket ignition modes. Then, constraints, such as terminal height, velocity, flight path angle, and angle of attack, are converted into the constraints within height-time profile according to the second-stage rocket flight characteristics. The closed-loop guidance method is inferred by different spline curves given the different terminal constraints. Afterwards, a thrust bias energy management strategy is proposed to waste the excess energy of the solid rocket. Finally, the proposed method is verified through nominal and dispersion simulations. The simulation results show excellent applicability and robustness of this method, which can provide a valuable reference for the ascent guidance of solid rocket-powered launch vehicles.

Multicriteria motion optimization of automatically controlled aerostatic vehicle

An approach to the multicriteria motion optimization of multimode automatically controlled aerostaticvehicle is offered. The approach is based on the mathematical apparatus of differential transformationof functions and equations and scalar convolution of partial criteria under the nonlinear scheme ofcompromises. Simulation of ascent trajectory using synthesized thrust vector control algorithm andwith taking into account the minimization of energy consumption and achieving of maximumhorizontal velocity at the end of the vehicle ascent has been conducted

Energy angular momentum closed-loop guidance

A novel guidance algorithm for launch vehicle ascent to the desired mission orbit is proposed. The algorithm uses total specific energy and orbital angular momentum as new state vector parameters. These parameters are ideally suited for the ascent guidance task, since the guidance algorithm steers the launch vehicle along a pre-flight optimal trajectory in energy angular momentum space. The guidance algorithm targets apogee, perigee, inclination and right ascension of ascending node. Computational complexities are avoided by eliminating time in the guidance computation and replacing it with angular momentum magnitude. As a result, vehicle acceleration, mass, thrust, length of motor burns, and staging times are also eliminated from the pitch plane guidance calculations. The algorithm does not involve launch vehicle or target state propagation, which results in minimal computational effort. Proof of concept of the new algorithm is presented using several numerical examples that illustrate performance results.

Ares I-X Trajectory Reconstruction: Methodology and Results

The Ares I-X trajectory reconstruction produced best-estimated trajectories of the flight-test vehicle ascent through stage separation and of the first- and upper-stage entries after separation. The trajectory-reconstruction process combines onboard, ground-based, and atmospheric measurements to produce the trajectory estimates, using an iterated extended Kalman filter algorithm. The Ares I-X vehicle had a number of onboard and ground-based sensors that were available, including inertial measurement units, radar, air data, and weather balloons. However, due to problems with calibrations and/or data, not all of the sensor data were used. This paper describes the methodology and results of the trajectory-reconstruction process, including flight-data preprocessing and input uncertainties, trajectory-estimation algorithms and dynamic models, output transformations, and comparisons with preflight predictions. The results of the reconstruction indicate nominal vehicle performance that is well within the range of expected dispersions based on preflight Monte Carlo analysis.

Lane changing patterns of bane and benefit: Observations of an uphill expressway

A mechanism is unveiled by which congestion forms on a 3-lane, uphill expressway segment, and causes reductions in output flow. Vehicular lane-changing (LC) is key to the mechanism, particularly LC induced by speed disturbances (SDs) that periodically arise in the expressway’s median and center lanes. Early in the rush, when flow was relatively low in the shoulder lane, drivers readily migrated toward that lane to escape the oncoming SDs. The shoulder lane thus acted as a ‘release valve’ for the high vehicular accumulations created by the SDs, such that forced vehicular decelerations were short-lived. The release valve failed only later in the rush, when flow increased in the shoulder lane in response to rising demand. LC induced by the SDs thereafter became disruptive: the decelerations they imposed spread laterally, and a persistent queue formed in all lanes. Long-run output flow dropped each day by 4–11% once the queue engulfed the base of the incline, and impeded vehicle ascent. Subtle details of this mechanism became visible by examining thousands of vehicle trajectories that were extracted from a series of eleven roadside video cameras. Though these trajectories were collected from only a single day, we suspect that the findings can be generalized to other days at the present site, and to other sites. This is because: (i) conspicuous features of the mechanism were repeatedly observed in loop detector data that were measured over many days at the site; (ii) these macroscopic features are consistent with observations previously made at other sites; and (iii) the more subtle details unveiled by the trajectories are compatible with a general theory of multi-lane traffic.

NASA Glenn Research Center Creek Road Complex—Cryogenic Testing Facilities

Due to expansion at neighboring Cleveland Hopkins Airport, several NASA Glenn Research Center (GRC) facilities have been relocated to the Creek Road Complex. The complex consists of the Small Scale Multi-purpose Research Facility (SMiRF), Cryogenic Components Lab Cell 7 (CCL-7), and a shop building. The facilities have been updated and include state-of-the art technology. SMiRF is a liquid hydrogen/liquid nitrogen (LH 2/LN 2) test facility used to conduct research in a 7400 L vacuum chamber. The chamber simulates space environment and launch vehicle ascent profile. SMiRF handles 5680 L of LH 2. CCL is a LH 2/LN 2 facility to perform small scale proof of concept tests for components and processes. It handles 1130 L of liquid hydrogen. Both facilities handle cryogens at sub-atmospheric pressures.

Optimization of Launch Vehicle Ascent Trajectories with Path Constraints and Coast Arcs

This paper describes improvements made to a hybrid analytic/numerical algorithm for optimization of launch vehicle trajectories. Modifications are described which improve the accuracy of the solution. In addition, the algorithm has been extended to include path constraints for normal force and angle of attack, and the terminal constraints have been generalized to allow optimal attachment to a target orbit defined by inclination, apogee radius and perigee radius. Singularities that occur for circular orbits and for equatorial orbits are identified. Finally, necessary conditions are derived and applied for the introduction of coasting arcs that are typically required for a great variety of missions.

Effect of Fence on Linear Aerospike Plume-Induced Base-Heating Physics

A computational heat transfer analysis is conducted to study the effect of the presence of an engine factor on the X-33 linear aerospike plume-induced base heating during vehicle ascent. A fence is an extension of plug side wall that reduces the plume spillage. A detailed three-dimensional thermoflowfield of the entire vehicle is computed such that an accurate freestream flow environment is assured for the base-flow development. The compupational methodology is based on a finite difference, viscous flow, chemically reacting, pressure-based computational fluid dynamics formulation and a finite volume, spectral-line-based weighted sum of gray gases absorption computational radiation heat transfer formulation. The computed base-flow physics are analyzed and compared with those of a subscale model hot-flow test. The effect of base bleed is also analyzed

A nonlinear programming approach for optimizing two-stage lifting vehicle ascent to orbit

An optimal atmospheric flight branched trajectory-shaping capability is presented based on the Davidon-Fletcher-Powell variable metric parameter optimization technique. Gradient information is generated using finite difference methods. A typical atmospheric flight branched optimization problem is analyzed which requires the determination of 31 parameters. This parameter set includes the three-dimensional description of vehicle attitude control angles for three branches of flight: first-stage ascent, second-stage ascent, and first- stage flyback. The important inflight inequality contraints required to maintain the integrity of the vehicles are considered. Some of the numerical methods employed are discussed, along with several new auxiliary techniques developed to improve the compatibility of the numerical gradient and iterator.