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NASA NTRS1991

Sensitivity of Space Station alpha joint robust controller to structural modal parameter variations

The photovoltaic array sun tracking control system of Space Station Freedom is described. A synthesis procedure for determining optimized values of the design variables of the control system is developed using a constrained optimization technique. The synthesis is performed to provide a given level of stability margin, to achieve the most responsive tracking performance, and to meet other design requirements. Performance of the baseline design, which is synthesized using predicted structural characteristics, is discussed and the sensitivity of the stability margin is examined for variations of the frequencies, mode shapes and damping ratios of dominant structural modes. The design provides enough robustness to tolerate a sizeable error in the predicted modal parameters. A study was made of the sensitivity of performance indicators as the modal parameters of the dominant modes vary. The design variables are resynthesized for varying modal parameters in order to achieve the most responsive tracking performance while satisfying the design requirements. This procedure of reoptimization design parameters would be useful in improving the control system performance if accurate model data are provided.

Kumar, Renjith R.↗

NASA NTRS2009

Deep-Space Ka-Band Link Priority Data Protection: Pre-Emptive Retransmission vs. Margin

In this paper the performance of two preemptive retransmission schemes for protection of priority data over deep-space Ka-band links is evaluated. The first scheme merges the correctly received bit from each transmission to create the most complete set of priority data for each pass (bit merge). The second scheme (symbol combining) combines the soft symbols received from each transmission of the priority data to increase the priority data's signal to noise ratio (SNR), thus increasing the liklihood of the correct reception.

Shambayati, Shervin↗

NASA NTRS2002

Centrifugal Compressor Surge Margin Improved With Diffuser Hub Surface Air Injection

Aerodynamic stability is an important parameter in the design of compressors for aircraft gas turbine engines. Compression system instabilities can cause compressor surge, which may lead to the loss of an aircraft. As a result, engine designers include a margin of safety between the operating line of the engine and the stability limit line of the compressor. The margin of safety is typically referred to as "surge margin." Achieving the highest possible level of surge margin while meeting design point performance objectives is the goal of the compressor designer. However, performance goals often must be compromised in order to achieve adequate levels of surge margin. Techniques to improve surge margin will permit more aggressive compressor designs. Centrifugal compressor surge margin improvement was demonstrated at the NASA Glenn Research Center by injecting air into the vaned diffuser of a 4:1-pressure-ratio centrifugal compressor. Tests were performed using injector nozzles located on the diffuser hub surface of a vane-island diffuser in the vaneless region between the impeller trailing edge and the diffuser-vane leading edge. The nozzle flow path and discharge shape were designed to produce an air stream that remained tangent to the hub surface as it traveled into the diffuser passage. Injector nozzles were located near the leading edge of 23 of the 24 diffuser vanes. One passage did not contain an injector so that instrumentation located in that passage would be preserved. Several orientations of the injected stream relative to the diffuser vane leading edge were tested over a range of injected flow rates. Only steady flow (nonpulsed) air injection was tested. At 100 percent of the design speed, a 15-percent improvement in the baseline surge margin was achieved with a nozzle orientation that produced a jet that was bisected by the diffuser vane leading edge. Other orientations also improved the baseline surge margin. Tests were conducted at speeds below the design speed, and similar results were obtained. In most cases, the greatest improvement in surge margin occurred at fairly low levels of injected flow rate. Externally supplied injection air was used in these experiments. However, the injected flow rates that provided the greatest benefit could be produced using injection air that is recirculating between the diffuser discharge and nozzles located in the diffuser vaneless region. Future experiments will evaluate the effectiveness of recirculating air injection.

Skoch, Gary J.↗

NASA NTRS2016

James Webb Space Telescope Deployment Brushless DC Motor Characteristics Analysis

A DC motor's performance is usually characterized by a series of tests, which are conducted by pass/fail criteria. In most cases, these tests are adequate to address the performance characteristics under environmental and loading effects with some uncertainties and decent power/torque margins. However, if the motor performance requirement is very stringent, a better understanding of the motor characteristics is required. The purpose of this paper is to establish a standard way to extract the torque components of the brushless motor and gear box characteristics of a high gear ratio geared motor from the composite geared motor testing and motor parameter measurement. These torque components include motor magnetic detent torque, Coulomb torque, viscous torque, windage torque, and gear tooth sliding torque. The Aerospace Corp bearing torque model and MPB torque models are used to predict the Coulomb torque of the motor rotor bearings and to model the viscous components. Gear tooth sliding friction torque is derived from the dynamo geared motor test data. With these torque data, the geared motor mechanical efficiency can be estimated and provide the overall performance of the geared motor versus several motor operating parameters such as speed, temperature, applied current, and transmitted power.

Tran, Ahn N.↗

NASA NTRS1998

An Adaptable Power System with Software Control Algorithm

A low cost, flexible and modular spacecraft power system design was developed in response to a call for an architecture that could accommodate multiple missions in the small to medium load range. Three upcoming satellites will use this design, with one launch date in 1999 and two in the year 2000. The design consists of modular hardware that can be scaled up or down, without additional cost, to suit missions in the 200 to 600 Watt orbital average load range. The design will be applied to satellite orbits that are circular, polar elliptical and a libration point orbit. Mission unique adaptations are accomplished in software and firmware. In designing this advanced, adaptable power system, the major goals were reduction in weight volume and cost. This power system design represents reductions in weight of 78 percent, volume of 86 percent and cost of 65 percent from previous comparable systems. The efforts to miniaturize the electronics without sacrificing performance has created streamlined power electronics with control functions residing in the system microprocessor. The power system design can handle any battery size up to 50 Amp-hour and any battery technology. The three current implementations will use both nickel cadmium and nickel hydrogen batteries ranging in size from 21 to 50 Amp-hours. Multiple batteries can be used by adding another battery module. Any solar cell technology can be used and various array layouts can be incorporated with no change in Power System Electronics (PSE) hardware. Other features of the design are the standardized interfaces between cards and subsystems and immunity to radiation effects up to 30 krad Total Ionizing Dose (TID) and 35 Mev/cm(exp 2)-kg for Single Event Effects (SEE). The control algorithm for the power system resides in a radiation-hardened microprocessor. A table driven software design allows for flexibility in mission specific requirements. By storing critical power system constants in memory, modifying the system code for other programs is simple. These constants can be altered also by ground command, or in response to an anomolous event. All critical power system functions have backup hardware functions to prevent a software or computer glitch from propagating. A number of battery charge control schemes can be implemented by selecting the proper control terms in the code. The architecture allows the design engineer to tune the system response to various system components and anticipated load profiles without costly alterations. A design trade was made with the size, weight and power dissipation of the electronics versus the performance of the power bus to load variations. Linear, fine control is maintained with a streamlined electronics design. This paper describes the hardware design as well as the software control algorithm. The challenges of closing the system control loop digitally is discussed. Control loop margin and power system performance is presented. Lab measurements are shown and compared to the system response of a hardware model running actual flight software.

Castell, Karen↗

NASA NTRS1993

Microscale diversity in satellite communications

Results from an investigation into a microscale diversity scheme for the purpose of mitigating attenuation due to tropospheric scintillations are presented. Data from an experiment using the 20 GHz Olympus satellite beacon were analyzed to evaluate the performance of short-baseline site-diversity systems. Results indicate that with a baseline separation of about 50 m the signals received by the main and diversity terminals become decorrelated. Examination of several high-scintillation events shows that such a diversity technique could substantially increase the uptime of low-power-margin VSAT systems. The performance of microscale diversity for several baseline separations is evaluated in terms of a parameter called diversity recovery. Applications and conditions under which microscale diversity performs best are addressed.

Cardoso, J. C.↗

NASA NTRS1970

Experimental evaluation of a honeycomb rotor shroud configuration to improve the stall margin of a 0.5 hub-tip ratio single-stage compressor. Volume 1 - Data and performance report

Experimental data and performance of honeycomb rotor shroud configuration to improve stall margin of 0.5 hub tip ratio single stage compressor

Oscarson, R. P.↗

NASA NTRS2025

Shock-Stationary Application of Pseudoshock Models During High-Amplitude Combustion-Driven Unsteadiness

The isolator pseudo-shock provides necessary compression within a dual-mode scramjet engine and buffers the engine system against unstart. Quasi-1D flux-conserved models are the state-of-the-art reduced-order model for optimization and online control of dual-mode scramjet engines. The stability and efficacy of this modeling approach is evaluated against data from a combustor-driven direct-connect experiment. The experiment exhibited strong combustor-driven unsteadiness that produced upstream propagating weak shocks into the isolator, interacting with the pseudo-shock. While this configuration resulted in unsteadiness that is atypical of standard operation, the experiment provided an opportunity to evaluate the modeling techniques in highly transient states. Such transients could occur during maneuvering or result from unexpected combustor events. A flexible quasi-1D formulation, the Fievet flux-conserved model is fit using Bayesian inference in the laboratory and shock-stationary reference frames. Model performance is analyzed using the Bayesian posteriors and model evaluations over the measured shock train speed range. It is concluded that to produce consistent isolator pressure profile estimates in this unsteady environment, the model must be implemented in a shock-stationary reference frame. Implementing this conclusion in model-based engine controllers may reduce needed unstart safety margins and increase maximum performance.

Bayesian↗

NASA NTRS2007

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗

NASA NTRS2006

Mars Science Laboratory: Entry, Descent, and Landing System Performance

Way, David W.↗

NASA NTRS2010

Use of ILTV Control Laws for LaNCETS Flight Research

A report discusses the Lift and Nozzle Change Effects on Tail Shock (LaNCETS) test to investigate the effects of lift distribution and nozzle-area ratio changes on tail shock strength of an F-15 aircraft. Specific research objectives are to obtain inflight shock strength for multiple combinations of nozzle-area ratio and lift distribution; compare results with preflight prediction tools; and update predictive tools with flight results. The objectives from a stability and control perspective are to ensure adequate aircraft stability for the changes in lift distribution and plume shape, and ensure manageable transient from engaging and disengaging the ILTV research control laws. In order to change the lift distribution and plume shape of the F-15 aircraft, a decade-old Inner Loop Thrust Vectoring (ILTV) research control law was used. Flight envelope expansion was performed for the test configuration and flight conditions prior to the probing test points. The approach for achieving the research objectives was to utilize the unique capabilities of NASA's NF-15B-837 aircraft to allow the adjustment of the nozzle-area ratio and/or canard positions by engaging the ILTV research control laws. The ILTV control laws provide the ability to add trim command biases to canard positions, nozzle area ratios, and thrust vectoring through the use of datasets. Datasets consist of programmed test inputs (PTIs) that define trims to change the nozzle-area ratio and/or canard positions. The trims are applied as increments to the normally commanded positions. A LaNCETS non-linear, six-degrees-of-freedom simulation capable of realtime pilot-in-the-loop, hardware-in-the-loop, and non-real-time batch support was developed and validated. Prior to first flight, extensive simulation analyses were performed to show adequate stability margins with the changes in lift distribution and plume shape. Additionally, engagement/disengagement transient analysis was also performed to show manageable transients.

Moua, Cheng↗

NASA NTRS1987

Ion accelerator system mounting design and operating characteristics for a 5 kW 30-cm xenon ion engine

Results from a series of experiments to determine the effect of accelerator grid mount geometry on the performance of the J-series ion optics assembly are described. Three mounting schemes, two flexible and one rigid, are compared for their relative ion extraction capability over a range of total accelerating voltages. The largest ion beam current, for the maximum total voltage investigated, is shown to occur using one of the flexible grid mounting geometries. However, at lower total voltages and reduced engine input power levels, the original rigid J-series ion optics accelerator grid mounts result in marginally better grid system performance at the same cold interelectrode gap.

Aston, Graeme↗

NASA NTRS2004

Space-Based Telemetry And Range Safety Flight Demonstration #1

The basic ability of STARS to maintain a satellite communications link with TDRSS satellites during dynamic aircraft flights was successfully demonstrated during FD 1. The Range Safety and Range User systems' link margins were measured. The ability to acquire/reacquire and maintain lock between a high-dynamic vehicle and a satellite-based system was demonstrated. The Range Safety system simultaneously received and processed command links from space and ground transmitters and provided near real-time Range Safety telemetry to DFRC, which then sent it in near real time to KSC, GSFC, and WFF for monitoring. The GPS receiver maintained track except during extremely dynamic maneuvers. The Range User system sent data at three different data rates. There were excellent cooperation and support from the different Centers, contractors, and Ranges. A large amount of data was recorded and extensive post-flight analysis was performed. The Range User TDRSS link margin met or exceeded the predicted performance at three different data rates. The Range Safety launch-head link margins generally agreed with the predicted performance. The UPS positions and velocities agreed with those from tracking radar to within about 20 m and a few rn/s. The link margins for the Range Safety TDRSS telemetry link were less than expected. The link margin for one TDRSS command link LPT channel was occasionally much less than the other. Additional post-flight testing has yet to identify the root causes of these results. There were many lessons learned from this first set of test flights. The most important one is that more time and testing are needed for each step to deal with the inevitable problems. It is vital that these lessons be among the primary areas of study that will carry over from FD#1 to FD#2, which is currently scheduled for early FY05 at DFRC and will use a specially designed Ku-band phased array antenna for the Range User system. The next series of flight demonstrations scheduled for late 2004 at DFRC will incorporate many lessons learned from FD#1. A specially designed Ku-band phased array antenna will be used with the Range User system. A test flight on a hypersonic vehicle is planned by the end of 2006.

Demspm. Erol↗

NASA NTRS1995

Dissipative rendering and neural network control system design

Model-based control system designs are limited by the accuracy of the models of the plant, plant uncertainty, and exogenous signals. Although better models can be obtained with system identification, the models and control designs still have limitations. One approach to reduce the dependency on particular models is to design a set of compensators that will guarantee robust stability to a set of plants. Optimization over the compensator parameters can then be used to get the desired performance. Conservativeness of this approach can be reduced by integrating fundamental properties of the plant models. This is the approach of dissipative control design. Dissipative control designs are based on several variations of the Passivity Theorem, which have been proven for nonlinear/linear and continuous-time/discrete-time systems. These theorems depend not on a specific model of a plant, but on its general dissipative properties. Dissipative control design has found wide applicability in flexible space structures and robotic systems that can be configured to be dissipative. Currently, there is ongoing research to improve the performance of dissipative control designs. For aircraft systems that are not dissipative active control may be used to make them dissipative and then a dissipative control design technique can be used. It is also possible that rendering a system dissipative and dissipative control design may be combined into one step. Furthermore, the transformation of a non-dissipative system to dissipative can be done robustly. One sequential design procedure for finite dimensional linear time-invariant systems has been developed. For nonlinear plants that cannot be controlled adequately with a single linear controller, model-based techniques have additional problems. Nonlinear system identification is still a research topic. Lacking analytical models for model-based design, artificial neural network algorithms have recently received considerable attention. Using their universal approximation property, neural networks have been introduced into nonlinear control designs in several ways. Unfortunately, little work has appeared that analyzes neural network control systems and establishes margins for stability and performance. One approach for this analysis is to set up neural network control systems in the framework presented above. For example, one neural network could be used to render a system to be dissipative, a second strictly dissipative neural network controller could be used to guarantee robust stability.

Gonzalez, Oscar R.↗

NASA NTRS1998

Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

The Liquid Fly-Back Booster (LFBB) Program seeks to enhance the Space Shuttle system safety, performance and economy of operations through the use of an advanced, liquid propellant Booster Main Engine (BME). There are several viable BME candidates that could be suitable for this application. The objective of this study was to identify the key Criteria to be applied in selecting among these BME candidates. This study involved an assessment of influences on the overall LFBB utility due to variations in the candidate rocket-engines characteristics. This includes BME impacts on vehicle system weight, performance, design approaches, abort modes, margins of safety, engine-out operations, and maintenance and support concepts. Systems engineering analyses and trade studies were performed to identify the LFBB system level sensitivities to a wide variety of BME related parameters. This presentation summarizes these trade studies and the resulting findings of the LFBB design teams regarding the BME characteristics that most significantly affect the LFBB system. The resulting BME choice should offer the best combination of reliability, performance, reusability, robustness, cost, and risk for the LFBB program.

Ryan, Richard M.↗

NASA NTRS2019

CPAS Airdrop Test Joint Recertification by Analysis

Airdrop testing of parachutes is a complicated endeavor that requires the custom design and certification of many critical components. The most direct path to certifying a component is to perform full scale testing with margin over the maximum loads expected to be seen in operation. However, other constraints often preclude the opportunity to perform full scale testing. In this paper, we present a case study where a problem arises in a joint that had been certified with a full scale test. There was no time or budget available to repeat the full scale testing after a redesign of the joint. Instead, we present a method of testing each failure mode at the component level to support a certification by analysis approach. The analysis itself was not complicated, but tradeoffs had to be made between different failure modes to arrive at the optimal design. The same approach was also applied back to the original joint to confirm that the failure mode that was not seen in full scale testing would have been caught by the proposed analysis. In the end, the new design was certified by analysis and worked without issue for the final six airdrop tests that used this joint.

McCann, Charles J.↗

NASA NTRS1972

Effect of commercial and military performance requirements for transport category aircraft on space shuttle booster design and operation

The effect of two sets of performance requirements, commercial and military, on the design and operation of the space shuttle booster is evaluated. Critical thrust levels are established according to both sets of operating rules for the takeoff, cruise, and go-around flight modes, and the effect on engine requirements determined. Both flyback and ferry operations are considered. The impact of landing rules on potential shuttle flyback and ferry bases is evaluated. Factors affecting reserves are discussed, including winds, temperature, and nonstandard flight operations. Finally, a recommended set of operating rules is proposed for both flyback and ferry operations that allows adequate performance capability and safety margins without compromising design requirements for either flight phase.

Bithell, R. A.↗

NASA NTRS1975

Space shuttle orbital maneuvering engine platelet injector program

A platelet-face injector for the fully reusable orbit maneuvering system OMS on the space shuttle was evaluated as a means of obtaining additional design margin and low cost. Performance, heat transfer, and combustion stability were evaluated over the anticipated range of OMS operating conditions. The effects of acoustic cavity configuration on combustion stability, including cavity depth, open area, inlet contour, and other parameters, were investigated using sea level bomb tests. Prototype injector and chamber behavior was evaluated for a variety of conditions; these tests examined the effects of film cooling, helium saturated propellants, chamber length, inlet conditions, and operating point, on performance, heat transfer and engine transient behavior. Helium bubble ingestion into both propellant circuits was investigated, as was chugging at low pressure operation, and hot and cold engine restart with and without a purge.

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