Robert D. Braun

Implementation and Performance Issues in Collaborative Optimization

Collaborative optimization is a multidisciplinary design architecture that is well-suited to large-scale multidisciplinary optimization problems. This paper compares this approach with other architectures, examines the details of the formulation, and some aspects of its performance. A particular version of the architecture is proposed to better accommodate the occurrence of multiple feasible regions. The use of system level inequality constraints is shown to increase the convergence rate. A series of simple test problems, demonstrated to challenge related optimization architectures, is successfully solved with collaborative optimization.

Collaborative Approach to Launch Vehicle Design

Collaborative optimization is a new design architecture specie cally created for large-scale distributed-analysis applications. In this approach, a problem is decomposed into a user-dee ned number of subspace optimization problems that are driven toward interdisciplinary compatibility and the appropriate solution by a system-level coordination process. This decentralized design strategy allows domain-specie c issues to be accommodated by disciplinary analysts while requiring interdisciplinary decisions to be reached by consensus. This investigation focuses on application of the collaborative optimization architecture to the multidisciplinary design of a singlestage-to-orbit launch vehicle. Vehicle design, trajectory, and cost issues are directly modeled in this problem, which is characterized by 95 design variables and 16 constraints. Numerous collaborative solutions are obtained. Comparison of these solutions demonstrates the ine uencethat an a priori ascent-abort criterion has on the vehicle design and the distinction between minimum weight and minimum cost concepts. The operational advantages of the collaborative optimization architecture include minimal framework integration requirements, the ability to use domain-specie c analyses, which already provide optimization without modie cation, inherent system e exibility and modularity, a distributed analysis and optimization capability, and a signie cant reduction in interdisciplinary communication requirements.

Use of the Collaborative Optimization Architecture for Launch Vehicle Design

Collaborative optimization is a new design architecture specifically created for large-scale distributed-analysis applications. In this approach, a problem is decomposed into a user-defined number of subspace optimization problems that are driven towards interdisciplinary compatibility and the appropriate solution by a system-level coordination process. This decentralized design strategy allows domain-specific issues to be accommodated by disciplinary analysts, while requiring interdisciplinary decisions to be reached by consensus. The present investigation focuses on application of the collaborative optimization architecutre to the multidisciplinary design of a single-stage-to-orbit launch vehicle. Vehicle design, trajectory, and cost issues are directly modeled. Posed to suit the collaborative architecture, the design problem is characterized by 95 design variables and 16 constraints. Numerous collaborative solutions are obtained. Comparison of these solutions demonstrates the influence which an a priori ascent-abort criterion has on development cost. Similarly, objective-function selection is discussed, demonstrating the difference between minimum weight and minimum cost concepts. The operational advantages of the collaborative optimization architecutre in a multidisciplinary design environment are also discussed.

Mars exploration entry, descent and landing challenges

The United States has successfully landed five robotic systems on the surface of Mars. These systems all had landed mass below 0.6 metric tons (t), had landed footprints on the order of hundreds of km and landed at sites below -1 km MOLA elevation due the need to perform entry, descent and landing operations in an environment with sufficient atmospheric density. Current plans for human exploration of Mars call for the landing of 40-80 t surface elements at scientifically interesting locations within close proximity (10s of m) of pre-positioned robotic assets. This paper summarizes past successful entry, descent and landing systems and approaches being developed by the robotic Mars exploration program to increased landed performance (mass, accuracy and surface elevation). In addition, the entry, descent and landing sequence for a human exploration system will be reviewed, highlighting the technology and systems advances required

Mars Pathfinder six-degree-of-freedom entry analysis

The Mars Pathfinder mission provides the next opportunity for scientific exploration of the surface of Mars following a 7.6 km/s direct entry. In support of this effort, a six-degree-of-freedom trajectory analysis and aerodynamic characteristic assessment are performed to demonstrate vehicle flyability and to quantify the effect that each of numerous uncertainties has upon the nominal mission profile. The entry vehicle is shown to be aerodynamically stable over a large portion of its atmospheric flight. Two low angle-of-attack static instabilities (freestream velocities of about 6.5 and 3.5 km/s) and a low angle-of-attack dynamic instability (supersonic) are identified and shown to cause bounded increases in vehicle attitude. The effects of center-of-gravity placement, entry attitude, vehicle roll rate, aerodynamic misprediction, and atmospheric uncertainty on the vehicle attitude profile and parachute deployment conditions are quantified. A Monte Carlo analysis is performed to statistically assess the combined impact of multiple off-nominal conditions on the nominal flight characteristics. These results suggest that there is a 99.7% probability that the peak attitude throughout the entry will be less than 8.5 deg, the peak heating attitude will be below 6.2 deg, and the attitude at parachute deployment will be less than 3.9 deg.

Influence of sonic-line location on Mars Pathfinder Probe aerothermodynamics

Flowfield solutions over the Mars Pathfinder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 deg are obtained. Aerodynamic coefficients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the movement of the sonic line between the blunted nose and the windside shoulder of the 70-deg half-angle cone in a gas with a low effective ratio of specific heats. The translation first occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40-km altitude, causing the effective specific heat ratio to decrease. The translation next occurs in an equilibrium flow regime as velocities decrease through 3.5 km/s and the specific heat ratio increases again with decreasing enthalpy. Laminar, windside heating levels may decrease with increasing angle of attack resulting from an increase in the effective radius of curvature with sonic line movement from the hemispherical nose to the aft shoulder of the blunt cone.

Mars Pathfinder atmospheric entry - Trajectory design and dispersion analysis

The Mars Pathfinder spacecraft will enter the Martian atmosphere directly from the interplanetary trajectory, with a velocity of up to 7.35 km/s. The definition of the nominal entry trajectory and the accurate determination of potential trajectory dispersions are necessary for the design of the Pathfinder entry, descent, and landing system. Monte Carlo numerical simulations have been developed to quantify the range of possible entry trajectories and attitude profiles. The entry trajectory requirements and constraints are discussed, and the design approach and uncertainties used in the Monte Carlo analysis are described. Three-degree-of-freedom and six-degree-offreedom trajectory results are compared. The Monte Carlo analysis shows that the Mars Pathfinder parachute will be deployed within the required ranges of dynamic pressure, Mach number, and altitude, over a 3<r range of trajectories. The Pathfinder 3<r landing ellipse is shown to be roughly 50 X 300 km. Nomenclature B - R = component of the 5-plane miss vector along the R axis, km

Guidance, Navigation, and Control System Performance Trades for Mars Pinpoint Landing

Landing site selection is a compromise between safety concerns associated with the site’s terrain and scientific interest. Therefore, technologies enabling pinpoint landing performance (sub-100-m accuracies) on the surface of Mars are of interest to increase the number of accessible sites for in situ research, as well as allow placement of vehicles nearby prepositioned assets. A survey of the performance of guidance, navigation, and control technologies that could allow pinpoint landing to occur at Mars was performed. This assessment has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at the end of the hypersonic guidance phase (parachute deployment) below approximately 3 km. Four different propulsive terminal descent guidancealgorithms were examined. Of these four, a near propellant-optimal analytic guidance law showed promisefortheconceptualdesignofpinpointlandingvehicles.Theexistenceofapropellantoptimumwithregardto theinitiationtimeofthepropulsiveterminaldescentwasshowntoexistforvarious flightconditions.Subsonicguided parachutes were shown to provide marginal performance benefits, due to the timeline associated with descent through the thin Mars atmosphere. This investigation also demonstrates that navigation is a limiting technology for Mars pinpoint landing, with landed performance being largely driven by navigation sensor and map tie accuracy.

Survey of Ballute Technology for Aerocapture

Ballute aerodynamic decelerators have been studied since early in the space age (1960’s), being proposed for aerocapture in the early 1980’s. Significant technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. The mass savings of this concept enables a number of high value science missions. Current studies of ballute aerocapture at Titan and Earth may lead to flight test of one or more ballute concepts within the next five years. This paper provides a survey of the literature with application to ballute aerocapture. Special attention is paid to advances in trajectory analysis, hypersonic aerothermodynamics, structural analysis, coupled analysis, and flight tests. Advances anticipated over the next 5 years are summarized.

High Mass Mars Entry, Descent, and Landing Architecture Assessment

As the nation sets its sight on returning humans to the Moon and going onward to Mars, landing high mass payloads ( 2 t) on the Mars surface becomes a critical technological need. Viking heritage technologies (e.g., 70 sphere-cone aeroshell, SLA-561V thermal protection system, and supersonic disk-gap-band parachutes) that have been the mainstay of the United States’ robotic Mars exploration program do not provide sucient capability to land such large payload masses. In this investigation, a parametric study of the Mars entry, descent, and landing design space has been conducted. Entry velocity, entry vehicle conguration, entry vehicle mass, and the approach to supersonic deceleration were varied. Particular focus is given to the entry vehicle shape and the supersonic deceleration technology trades. Slender bodied vehicles with a lift-to-drag ratio (L=D) of 0.68 are examined alongside blunt bodies with L=D = 0.30. Results demonstrated that while the increased L=D of a slender entry conguration allows for more favorable terminal descent staging conditions, the greater structural eciencies of blunt body systems along with the reduced acreage required for the thermal protection system aords an inherently lighter vehicle. The supersonic deceleration technology trade focuses on inatable aerodynamic decelerators (IAD) and supersonic retropropulsion, as supersonic parachute systems are shown to be excessively large for further consideration. While entry masses (the total mass at the top of the Mars atmosphere) between 20 and 100 t are considered, a maximum payload capability of 37.3 t results. Of particular note, as entry mass increases, the gain in payload mass diminishes. It is shown that blunt body vehicles provide sucient vertical L=D to decelerate all entry masses considered through the Mars atmosphere with adequate staging conditions for the propulsive terminal descent. A payload mass fraction penalty of approximately 0.3 exists for the use of slender bodied vehicles. Another observation of this investigation is that the increased aerothermal and aerodynamic loads induced from a direct entry trajectory (velocity 6.75 km/s) reduce the payload mass fraction by approximately 15% compared to entry from orbital velocity ( 4 km/s). It should be noted that while both IADs and supersonic retropropulsion were evaluated for each of the entry masses, congurations, and velocities, the IAD proved to be more mass-ecient in all instances. The sensitivity of these results to modeling assumptions was also examined. The payload mass of slender body vehicles was observed to be approximately four times more sensitive to modeling assumptions and uncertainty than blunt bodies.