Computer recommendations for an automatic approach and landing system for V/STOL aircraft. Volume 1 - Computer recommendations
Evaluation of digital computer for V/STOL aircraft automatic approach and landing system
Gaines, H. T.↗
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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Landing performance of an air cushion landing system installed on a 1/10-scale dynamic model on the C-8 Buffalo airplane
An experimental study was conducted to evaluate the landing behavior of a 1/10-scale dynamic model of the C-8 Buffalo airplane equipped with an air-cushion landing system (ACLS) on a variety of surfaces including both calm and rough water and a smooth hard surface. Taxi runs were made on the hard surface over several obstacles. Landings were made with the model at various pitch and roll attitudes and vertical velocities and at one nominal horizontal velocity. Data from the landings include time histories of the trunk and air-cushion pressures and accelerations at selected locations on the model.
Thompson, W. C.↗
Minimum variance estimates of signal derivatives - A problem in instrument landing systems
Kalman filter application as observer of observable signals derivatives, using gain matrix to minimize variance estimate for instrument landing systems
Martin, J. C.↗
Experimental determination of position-estimate accuracy using back-azimuth signals from a microwave landing system
Flight tests using the Boeing 737 airplane to obtain position estimates with back azimuth signals from a microwave landing system (MLS) are discussed. The equations and logic used to generate a navigation position estimate in the MLS back azimuth signal environment are described. The error in the navigation position estimate is determined. A summary of the Boeing 737 position estimate update process is described. The navigation position estimate error calculated flight data and radar tracking information is analyzed. The position estimate error data using the MLS inputs are compared with error data obtained during dual distance measuring equipment updates.
Knox, C. E.↗
Parametric Mass Modeling for Mars Entry, Descent and Landing System Analysis Study
This paper provides an overview of the parametric mass models used for the Entry, Descent, and Landing Systems Analysis study conducted by NASA in FY2009-2010. The study examined eight unique exploration class architectures that included elements such as a rigid mid-L/D aeroshell, a lifting hypersonic inflatable decelerator, a drag supersonic inflatable decelerator, a lifting supersonic inflatable decelerator implemented with a skirt, and subsonic/supersonic retro-propulsion. Parametric models used in this study relate the component mass to vehicle dimensions and mission key environmental parameters such as maximum deceleration and total heat load. The use of a parametric mass model allows the simultaneous optimization of trajectory and mass sizing parameters.
Samareh, Jamshid A.↗
Braking and cornering studies on an air cushion landing system
An experimental investigation was conducted to evaluate several concepts for braking and steering a vehicle equipped with an air cushion landing system (ACLS). The investigation made use of a modified airboat equipped with an ACLS. Braking concepts were characterized by the average deceleration of the vehicle. Reduced lobe flow and cavity venting braking concepts were evaluated in this program. The cavity venting braking concept demonstrated the best performance, producing decelerations on the test vehicle on the same order as moderate braking with conventional wheel brakes. Steering concepts were evaluated by recording the path taken while attempting to follow a prescribed maneuver. The steering concepts evaluated included using rudders only, using differential lobe flow, and using rudders combined with a lightly loaded, nonsteering center wheel. The latter concept proved to be the most accurate means of steering the vehicle on the ACLS, producing translational deviations two to three times higher than those from conventional nose-gear steering. However, this concept was still felt to provide reasonably precise steering control for the ACLS-equipped vehicle.
Daugherty, R. H.↗
Wireless Channel Characterization in the 5 GHz Microwave Landing System Extension Band for Airport Surface Areas
In this project final report, entitled "Wireless Channel Characterization in the 5 GHz Microwave Landing System Extension Band for Airport Surface Areas," we provide a detailed description and model representation for the wireless channel in the airport surface environment in this band. In this executive summary, we review report contents, describe the achieved objectives and major findings, and highlight significant conclusions and recommendations.
Matolak, David W.↗
What's Cookin' With Helicopter Microwave Landing Systems?
This article describes a joint effort of the FAA, NASA and the Helicopter Industry to establish a data base for (Microwave Landing Systems) MLS approaches, and to aid the FAA in developing the MLS Terminal Instrument Procedures (TERPS) criteria for rotary-wing IFR operations. The latter is of particular importance for it appears to signal sincere official interest in filling in the blanks in Chapter 11 of TERPS -- that portion set aside for helicopter IFR operations and recognition of the unique capabilities of the helicopter in the instrument environment. One of the program objectives is to find out what a sample of pilots can do with the MLS in a minimum machine.
Taylor, Richard L.↗
Human Landing System Storable Propellant Architecture: Mission Design, Guidance, Navigation, and Control
In preparation of the NextSTEP-2 Appendix H awards, NASA’s Human Landing System (HLS) Program (HLSP) went through an abbreviated Design Analysis Cycle (DAC), or miniDAC, to characterize and quantify how a storable propellant design would affect and improve overall system performance of the Government reference design. As part of that miniDAC, HLSP ran multiple trade studies and assessments, including aspects of Mission Design, Flight Mechanics (FM), and Guidance, Navigation, and Control (GN\&C). The storable propellant designs were assessed through the baseline mission targeting the Near Rectilinear Halo Orbit (NRHO) and several alternate staging orbits. This paper focuses on the FM and GN\&C trades and assessments completed as part of the miniDAC as well as the GN\&C elements of the miniDAC Master Equipment List (MEL) developed for the storable propellant lander configurations.
A S Craig↗
Human Landing System Storable Propellant Architecture: Mission Design, Guidance, Navigation, and Control
In preparation of the NextSTEP-2 Appendix H awards, NASA’s Human Landing System (HLS) Program (HLSP) went through an abbreviated Design Analysis Cycle (DAC), or miniDAC, to characterize and quantify how a storable propellant design would affect and improve overall system performance of the Government reference design. As part of that miniDAC, HLSP ran multiple trade studies and assessments, including aspects of Mission Design, Flight Mechanics (FM), and Guidance, Navigation, and Control (GN&C). The storable propellant designs were assessed through the baseline mission targeting the Near Rectilinear Halo Orbit (NRHO) and several alternate staging orbits. This paper focuses on the FM and GN&C trades and assessments completed as part of the miniDAC as well as the GN\&C elements of the miniDAC Master Equipment List (MEL) developed for the storable propellant lander configurations.
Landers↗
Research study of an aircraft-contained radar zero-zero landing system, volume 1 Final report, Jan. - Dec. 1967
Two rapid scan radar/display techniques evaluated for use in aircraft-contained zero-zero landing system
Suzansky, J. W.↗
Modeling of Instrument Landing System (ILS) localizer signal on runway 25L at Los Angeles International Airport
A joint NASA/FAA flight test has been made to record instrument landing system (ILS) localizer receiver signals for use in mathematically modeling the ILS localizer for future simulation studies and airplane flight tracking tasks. The flight test was conducted on a portion of the ILS localizer installed on runway 25L at the Los Angeles International Airport. The tests covered the range from 10 to 32 n.mi. from the localizer antenna. Precision radar tracking information was compared with the recorded localizer deviation data. Data analysis showed that the ILS signal centerline was offset to the left of runway centerline by 0.071 degrees and that no significant bends existed on the localizer beam. Suggested simulation models for the ILS localizer are formed from a statistical analysis.
Hueschen, Richard M.↗
Wireless Channel Characterization: Modeling the 5 GHz Microwave Landing System Extension Band for Future Airport Surface Communications
We describe a recently completed wideband wireless channel characterization project for the 5 GHz Microwave Landing System (MLS) extension band, for airport surface areas. This work included mobile measurements at large and small airports, and fixed point-to-point measurements. Mobile measurements were made via transmission from the air traffic control tower (ATCT), or from an airport field site (AFS), to a receiving ground vehicle on the airport surface. The point-to-point measurements were between ATCT and AFSs. Detailed statistical channel models were developed from all these measurements. Measured quantities include propagation path loss and power delay profiles, from which we obtain delay spreads, frequency domain correlation (coherence bandwidths), fading amplitude statistics, and channel parameter correlations. In this paper we review the project motivation, measurement coordination, and illustrate measurement results. Example channel modeling results for several propagation conditions are also provided, highlighting new findings.
Matolak, D. W.↗
Probabilistic Modeling of a Three-Stage Human Landing System Architecture
Space Policy Directive-1 has led to NASA partnerships with commercial entities on procurement which includes the development of the Human Landing System (HLS) [1]. With the goal of delivering human crew to the lunar surface by 2024, system uncertainties become an important obstacle to the maturation of multiple new, driving technologies and mission concepts of the HLS program. As unmitigated uncertainties have previously led to failed development programs, these risks and their impacts must be understood and handled to ensure program success [2]. Sources of uncertainty include novel engine designs and configurations, increased reliance on cryogenic fluid management(CFM), and refueling technologies—which propagate as high-level performance metrics such as overall propellant mass and engine performance. Also, the occurrence of operational uncertainties—e.g. launch conditions or need to abort during the mission—can cause cascading effects on the rest of the mission that are difficult to definitively quantify, and are outside the scope of control. These concrete examples and other occurrences can be categorized as either epistemic or aleatory uncertainties.Epistemic uncertainty arises due to a lack of knowledge and can be alleviated with design and program maturation. Aleatory uncertainty is due to the inherent randomness of the system and cannot be directly reduced, unlike epistemic uncertainty. Robust design and probabilistic methods can compensate for aleatory effects. A taxonomy of uncertainty is referred to for this work [3]. In this paper, a probabilistic methodology to handle uncertainties has been demonstrated on a three-element HLS concept [1, 4], which allows tracking of current best estimates of the concept and assessment of concept design robustness against uncertainties. A sample case has been completed for this abstract, and an expansion on the methodology will be included in the final paper. This methodology has two key parts: first, the creation of a dynamic architecture model of a three-element HLS concept; and second, its use with surrogate modeling and range estimating techniques to capture and propagate uncertainties. This abstract will cover the basics of the approach used, and further details and justifications will be in the final paper.The mission profile associated with this three-element concept (Fig 1) was modeled as a set of mission events that facilitated mass changes, idles, or spacecraft maneuvers. The mission profile scope starts with each element’s NRHO orbit insertion and aggregation and ends at post-sortie rendezvous with Orion. More detail on the mission profile will be in the final paper. The DYnamic Rocket EQuation Tool (DYREQT), a space systems synthesis and sizing framework used by NASA, was used as the physics framework to model the HLS architecture for applying the probabilistic methodology [5, 6]. Specifically, a parametric representation of the lander, ascent, and transfer elements and the mission profile of each element was established, with vehicle and mission parameters available as inputs to allow for a dynamic model. Each vehicle stage was modeled with high-level performance metrics, using Isp and propellant mass fraction (PMF) to remain parametric. For the probabilistic analysis, uncertainties of interest within the HLS concept were enumerated and represented as parameters within the DYREQT model as inputs for vehicle stages or mission profile events. These parameters were frozen at their nominal values for the purposes of baselining architecture performance and sizing the vehicle appropriately based on reference documentation [1]. Range estimating—a probabilistic method that combines Monte Carlo sampling, focus on critical parameters, and heuristics to assess risk and opportunities—is traditionally used with Mass Equipment Lists (MELs), but has been adapted with operational parameters as well as vehicle parameters in theDYREQT model to capture mission uncertainty alongside vehicle uncertainty [7, 3]. This method was selected due to its application and insight on a system from a bottom-up perspective, independence from historical rules of thumb, and ability to generate sensitivities based on design decisions and uncertainties. As a sample case for the abstract, the boiloff rates of the vehicle elements and the loiter times during the mission (simulating launch time variations and changing window of opportunities) were used with range estimating to provide preliminary results. To perform the range estimation portion of this methodology (depicted in Fig. 3, further details in final paper), the DYREQT model was sampled using a Design of Experiments (DoE) to efficiently explore the architecture design space with respect to the sample set of uncertainty parameters; 5,000 cases via Latin Hypercube Sampling were computed on the DYREQT architecture model. Then, the results were used to create surrogate models, multivariate regressions that can visualize hypercube trends in the design space, of the architecture with respect to the uncertainty parameters. Range estimating was applied to the surrogates instead of the actual models, which saves computational expense due to the bulk of cases needed for the Monte Carlo simulation as part of range estimating. Uncertainty parameters were sampled independently from triangular distributions using the DoE ranges as ‘min’ and ‘max’, and the nominal value as ‘most likely’. Based engineering intuition, some uncertainty parameters are correlated—e.g. if the main propellant has a high boil-off rate, the oxidizer should follow suit as both are related to CFM technology.While a Monte Carlo simulation samples all inputs as independent, the results would show model correlations; thus, it is efficient to sample the inputs as correlated. Using a correlation matrix constructed for the uncertainty parameters, previously independent samples were transformed to perform a Correlated Monte Carlo. A table for the DoE ranges and probability distribution parameters is shown in Table 1, and more details on Correlated Monte Carlo Simulations will be discussed in the final paper. The model’s resulting DoE showed that multivariate polynomial equations fit via least squares method captured its behavior accurately for the sample case. For the Correlated Monte Carlo Simulation, a positive correlation between fuel and oxidizer boiloff rates was used as a demonstration. 10,000 cases were computed with the surrogates and the launched masses for each vehicle element was collated. The results can be displayed in a probability density function (PDF), showing the impact of the uncertainty parameters chosen. Integrating the PDFs will yield a cumulative distribution function (CDF) that shows the cumulative probability of a given value on the x-axis. For the sample case, the elements’ launch mass margin was calculated and represented in as CDFs, as a demonstrated representation of figures of merit for the HLS concept. For the lander and ascent elements, the NRHO mass insertion limit is 16t; the transfer element has a limit of 30t [1]. It can be seen with Figure 2 that this probabilistic methodology can provide insight into mass margin with respect to the uncertainties being modeled. Currently, the results show that the lander (descent) vehicle element has the most restrictive design space; it is the only element to show a 10% probability of negative margin. Further analysis on the Monte Carlo results will show sensitivities for driving constraints and parameters for architecture feasibility, which can lead to establishing potential mission rules.The combination of range estimating with a parametric architecture model for HLS demonstrated the capability of this probabilistic methodology in a sample case. As the HLS development progresses, this methodology has the potential for keeping current best estimates of architecture performance for awarded concepts due to the flexibility in DYREQT’s modeling framework and its parametric nature. Concept maturation and increased epistemic knowledge can be injected into the model probabilistic modeling, and thus continue to track probability of mission success.
Stephanie Y Zhu↗
A Low-Cost and Low-Risk Testbed for Control Design of Launch Vehicles and Landing Systems
This research seeks to bridge the gap between the simulation/theory and the flight-testing phases of control algorithmdesign for launch vehicles (LVs) and landing systems (LSs). Thispaper reviews contemporary testbeds for LV and LS control sys-tem design with an emphasis on reducing costs and risks. Also, anew low-cost and low-risk testbed utilizing a quadcopter, flexibleinverted pendulum, and a hanging pendulum is proposed. It is shown that the planar dynamic response of this proposed testbedcan match the planar dynamic response of an LV. As mission objectives for frequent cislunar and interplanetary travel become more complex, the algorithms that control LVs and LSs need to become more advanced to ensure mission safety and success. These control systems are slow to mature because it is difficult to perform experiments in an environment relevant to the mission in a safe and cost-effective manner. A solution to this problem frequently leveraged in the aviation andsatellite industries involves mimicking the dynamic response ofthe vehicle with a test platform that is low-cost and low-risk.
William J Elke III↗
Enabling Global Lunar Access for Human Landing Systems Staged at Earth-Moon L2 Southern Near Rectilinear Halo and Butterfly Orbits
NASA’s Artemis program will return astronauts to the lunar surface using a Human Landing System staged at Gateway’s 9:2 L2 southern NRHO. The current study explores global lunar access for Gateway’s baseline NRHO and a comparable butterfly orbit. Additionally, increasing transfer time and loitering in low lunar orbit are analyzed to reduce lunar surface access costs. The results show selecting the optimal staging orbit for a landing site, increasing transfer times, and leveraging low lunar orbit loiter greatly reduce performance requirements.
Zachary D May↗
The Unparalleled Systems Engineering of MSL's Backup Entry, Descent, and Landing System: Second Chance
Second Chance (SECC) was a bare bones version of Mars Science Laboratory's (MSL) Entry Descent & Landing (EDL) flight software that ran on Curiosity's backup computer, which could have taken over swiftly in the event of a reset of Curiosity's prime computer, in order to land her safely on Mars. Without SECC, a reset of Curiosity's prime computer would have lead to catastrophic mission failure. Even though a reset of the prime computer never occurred, SECC had the important responsibility as EDL's guardian angel, and this responsibility would not have seen such success without unparalleled systems engineering. This paper will focus on the systems engineering behind SECC: Covering a brief overview of SECC's design, the intense schedule to use SECC as a backup system, the verification and validation of the system's "Do No Harm" mandate, the system's overall functional performance, and finally, its use on the fateful day of August 5th, 2012.
fault protection↗
Evaluation of microwave landing system approaches in a wide-body transport simulator
The objective of this study was to determine the suitability of flying complex curved approaches using the microwave landing system (MLS) with a wide-body transport aircraft. Fifty pilots in crews of two participated in the evaluation using a fixed-base simulator that emulated an MD-11 aircraft. Five approaches, consisting of one straight-in approach and four curved approaches, were flown by the pilots using a flight director. The test variables include the following: (1) manual and autothrottles; (2) wind direction; and (3) type of navigation display. The navigation display was either a map or a horizontal situation indicator (HSI). A complex wind that changed direction and speed with altitude, and included moderate turbulence, was used. Visibility conditions were Cat 1 or better. Subjective test data included pilot responses to questionnaires and pilot comments. Objective performance data included tracking accuracy, position error at decision height, and control activity. Results of the evaluation indicate that flying curved MLS approaches with a wide-body transport aircraft is operationally acceptable, depending upon the length of the final straight segment and the complexity of the approach.
Summers, L. G.↗