Aircraft Noise and Performance Data for a Notional Supersonic Business Jet

th 28 AIAA/CEAS Aeroacoustics Conference AIAA 2022 - XXXX Southampton, UK, 14 - 17 June, 2022

Aircraft Noise and Performance Data for a Notional

Supersonic Business Jet

* Jeffrey J. Berton NASA Glenn Resear ch Center, Cleveland, Ohio 44135 The International Civil Aviation Organization recently completed a study that determined the global environmenta l impact of adding several hypothetical supersonic aircraft types to the existing subsonic fleet. NASA support ed this study by designing a notio nal 55 - tonne supersonic business jet . The airplane is named the Supersonic Technology Concept Aeroplane by the i nternational community . Performance, noise, and exhaust emission predictions for this transport were used to inform development of new environme ntal standards for future supersonic civil aircraft. The behavior of the aircraft in an operational setting is c onsidered.

Calculation of aircraft noise and performance data for th is aircraft is the focus of this paper.

Noise abatement departure procedures , stage length performance, and noise - power - distance data are determined using NASA tools. Also investigated are some of the anticipated behaviors and requirements of supersonic aircraft in the commercial airspace.

I. Introduction † UPERSONIC Technology Concept Aeroplanes are conceptual research vehicles studied with th e intent to provide information to t he Internatio nal Ci vil Aviation Organization by way of their Committee on Aviation

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Environmental Protection (CAEP) . Though several STCAs are being investigated by CAEP, the focus of this paper is an analysis of an eight - passenger , Mach 1.4 business jet developed by NAS A . The airplane is intended to be generally representative of early market entrant supersonic business jet aircraft being considered b y industry. It is designed to travel transatlantic distances at Mach 1.4 using relatively near - term technologies. It is eq uipped with conceptual engines designed and assessed by NASA. They are derived from the core of a contemporary, “off - the - shelf” subson ic turbofan . All aspects of the airplane and engine have been developed entirely with information from the public domain.

Being nonproprietary and transparent, the STCA is ideal for use in ICAO’s public studies. NASA began work on the STCA in 2017 in suppo rt of ICAO. Since then, the STCA continues to serve as a notional reference airplane used in studies conducted by NASA [ 1 - 3 ] and by others [ 4 - 11 ]. Aircraft mission performance, airport - vicinity noise, and exhaust emissions are predicted for the STCA using NASA tools.

There are three motivations for this study. First, during the eleventh triennial cycle of CAEP, two subgroups were formed in 2016 by the r a pporteurs of CAEP’s noise and exhaust emissions working groups. The charter of these subgroups is to examine the suitability of existing noise and exhaust emission cert ification standards for new supersonic civil airplane types. The subgroups are tasked with developing appropriate standards and recommended practices to support anticipated entries of n ew supersonic civil transports. The ultimate goal of both subgroups is to recommend amendments to Part II , Chapter 12 of [ 12 ] and to Part III, Chapter 3 of [ 13 ] that more properly address noise and exhaust emissions of supersonic airplane types, respective ly. Both subgroups are dedicated to ensuring that new supersonic transpor ts operate responsibly in an environmentally sustainable manner. NASA is assisting these subgroups by acting as analysts and independent arbitrators and by providing unbiased, open as sessments of supersonic civil aircraft.

Second, the Federal Aviation Admi nistration Reauthorization Act of 2018 [ 14 ] requires the FAA to “exercise leadership in the creation of Federal and international policies, regulations, and standards relating to the certification and safe and efficient operation of civil supersonic aircraf t.” To this end, the FAA issue d a notice of proposed rulemaking regarding landing and takeoff noise standards for supersonic aircraft i n April, 2020 [ 15 ] . The FAA’s intention is to r evise its noise standards [ 16 ] to include provisions for future supersonic t ransports. NASA assisted the FAA in this effort.

* Aerospace Engineer, Propulsion Systems Analysis Branch , senior member AIAA.

† In referring to STCAs, “aeroplane” is the spelling used by ICAO. The spelling “airplane,” however, is used in this paper.

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

28th AIAA/CEAS Aeroacoustics Conference Third, at CAEP’s eleventh triennial meeting in 2019, the committee agreed to conduct an exploratory study on the global impact of f uture civil supersonic aircraft [ 17 ] . The FAA endorsed the study [ 18 ] and requ ested that NASA assist it by providing the STCA’s performance, fuel burn, noise , and exhaust emission data. The CAEP study predict ed the outcome of adding supersonic transports to the existing civil aircraft fleet. The study was part of CAEP’s twelfth (pre vious) program cycle and has since been completed. The STCA was an analytical proxy for future supersonic business jet types. The in fluence that supersonic airplanes are expected to have on operations , fuel consumption, airport noise, and air quality were assessed. Aircraft noise and performance data for the STCA were required to support the exploratory study. Derivation of the data is documented in this paper. A public report of the exploratory study is expected by the 2022 meeting of the CAEP Steering Gro up.

Presented in this paper is a description of the STCA’s propulsion and airframe analysis, operational n oise abatement departure p rocedures , stage length performance, and noise - power - distance data . Also investigated are some of the anticipated behaviors and requirements of these aircraft in the commercial airspace . A revised assessment of the STCA’s noise levels in a certification setting is described in a companion paper [ 19 ].

II. Airframe and Engines Prerequisites for computing aircraft noise and performance (ANP) data are working multidisciplinary models for the vehicle. Airframe, weight, aerodynamics, and propulsion an alyses taken from [ 2 ] are described in the following sections.

A. Concept Airframe The STCA is a notional, eight - pa ssenger business jet designed for Mach 1.4 supersonic overwater cruise speeds.

It has a low - aspect ratio cranked delta wing. Three engines are mounted aft; the outboard engines are mounted on short fuselage pylons, while the center engine is integrated wit h the vertical tail. A summary of vehicle characteris tics is shown in Table 1 . A solid model of the airplane is shown in Figure 1 (h alf of the top vi ew is drawn to show the interior arrangement and lo cations of fuel ).

The airplane has no features that would reduce its sonic boom noise signature, making it similar in that respect to other early market entrant designs. Supersonic speeds over land would be restricted where prohibited. Despite NASA’s current inter est in low - boom supersonic aircraft [ 20 ], those types of designs are viewed here as candidates for later market s .

Table 1 : STCA characteristics.

Max takeoff weight, klb 121 Passengers 8 Cruise Mach 1.4 Overall length, ft 135 Span, ft 67 Wing reference area, ft 1619 Wing aspect ratio 2.7 Wing taper ratio 0.09 Wing loading, lb/ft 74 Wing fuel, klb 24 Fuselage fuel, klb 36 Fu el fraction 0.50 Figure 1 . Solid model of the STCA .

Several aircraft conceptual design tools are used to synthesize the vehicle. A solid modeling tool [ 21 ] is used to define the outer mold lines of the airplane, to guide interior packaging, and to estimate internal fuel volume . The solid model also informs component weight and vehicle aerodynamic analyses. Lift - dependent drags, lift - independent drags, and wave drags are calculated by the methods described in Refs. [ 22 - 24 ]. A weight estimate of the win g is made using physics - based factors based on its gross geometry, while weight estimates of other major structures and systems are made using statistical - empirical relations. All airframe weight estimation methods are discussed in [ 25 ]. These computer code s are organized together using a frameworking tool [ 26 ] which provides a conceptual - level, multidisciplinary, integrated process for designing and analyzing supersonic aircraft [ 27 ]. Using this integrated design 2 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference environment, major aircr aft design variables c an be rapidly assessed and optimized. A design and sizing analysis of wing, fuselage and tail characteristics is performed, subject to practical performance constraints.

The vehicle design is optimized to maximize ra nge for a fixed ma ximum gross weight of 55t (121k lb ) . S ubject to a minimum cabin wi d th of 7ft , f uselage section height, width, and tangent angles at three stations are optimized to minimize wave dr ag. Wing planform shape, a irfoil twist , and camber are varied to minimize wav e drag and lift - depend ent drag. Horizontal tail size and location, main gear location , and trailing edge and leading edge flap deflections are optimized to meet takeoff, landing, and cruise static margin constraints , and to ensure reasona ble takeoff and la nding field lengths an d approach velocity.

B. Propulsion For an early market entrant, i t is unlikely that a completely new engine could be developed and be ready in time for a near - term entry into service . Instead, it is more likely that the low - pressure spool of a contemporary of f - the - shelf engine would be redesigned and repurposed , resulting in a supersonic variant of an existing subsonic turbofan. In this study, an analytical model of a subsonic CFM Int ernational CFM56 - 7B 27 is used as the “donor” engine from which the supersonic engine is derived. Interestingly, the original CFM56 - 2, granted certification in 1979, was itself derived from the General Electric F101 model used for the supersonic B - 1A bomber . R edesigning the low - pressure spool of a CFM56 once again for a supersonic a pplication would bring the engine family full circle .

Because much engine design data are closely - held, proprietary , and unavailable, any analytic simulation of a CFM56 (outside o f CFM International) will necessarily have some inherent inaccuracy. Neverthe less, if data can be obtained from public - domain sources (such as type certificate data sheets, manufacturer - provided operating documents, technical reports , and manufacturer’s we bsites), simulations of turbofans developed outside of engine companies can b e reasonably accurate. A model of the subsonic CFM56 is created with such information using the Numerical Propulsion System Simulation code (NPSS, [ 28 , 29 ]) to predict engine perf ormance. NPSS is an engine cycle analysis tool developed jointly by NASA and by United States aerospace industry. It is currently the accepted, state - of - the - art software for airbreathing engine cycle performance analysis for United States industry, academi a, and NASA. The subsonic CFM56 model is adapted from work performed under the FAA’s Environmental Design Space init iative [ 30 , 31 ].

The low - pressure spool of the CFM56 - 7B is redesigned for a Mach 1.4 cruise application. The booster is discarded (with it, supersonic ram effects would elevate aft stages of the compressor to excessive te mperature), and the fan and low - pressure turbine are redesigned for a higher pressure ratio. Fan performance is modeled using data collected at NASA from the GE57 scale mo del fan [ 32 ]. The GE57 fan is considered to be perhaps representative of what might be used by an engine manufacturer in a supersonic refan application. It consists of a single stage and operates at peak efficiency at a pressure ratio of 2.2. Fan pressure ratio is a design variable strongly influencing engine performance.

A high fan pressur e ratio is preferred to create an exhaust velocity high enough for supersonic flight, while a low fan pressure ratio is required to meet takeoff and landing noise require ments. Fan pressure ratio, along with a practical extraction ratio, directly determine the bypass ratio of the engine. This poses conflicting requirements for supersonic engine designers. If fan pressure ratios are high enough, they could lead to supercrit ical nozzle pressure ratios at low altitude and create high levels of jet shock cell n oise during takeoff.

Another design choice is whether to forcibly mix the core and bypass streams or to allow them to remain separate.

There are compelling reasons to mix the streams. There is usually an increase in gross thrust when flows are forcibly mix ed and exhausted through a common nozzle, with the benefit increasing with increasing core stream temperature.

And the outer mold lines of a simpler, single - stream nozzle are preferred over those of a more complex coannular nozzle if sonic boom reduction i s important. In this study, core and bypass streams are forcibly mixed through a lobed mixer. The design extraction ratio is kept near unity so that mixer bypass port and mixer exit Mach numbers are always less than 0.5.

The mixed flow exits through a sing le - stream convergent - divergent plug nozzle. The centerbody plug and nozzle throat are fixed while the divergent flaps are variable. The plug is important in keeping aftbo dy boattail angles small during supersonic cruise while also lowering takeoff jet nois e slightly. At low altitudes, the divergent nozzle flaps are closed to a minimum area so that the nozzle exit plane is the throat. Solid models of the CFM56 - 7B and the de rived supersonic variant are shown in Figure 2 . Not shown in the figure are the nozzles or inlets for either engine .

3 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference At the cycle design point, characteristics of the compressor and the high - pressure turbine are set manually to those of the CFM56 donor engine. This keeps the high - pressure spool of the supersonic derivative engine identical to the CFM56 core. Bleed f low fractions and core flow passage areas are also held constant. Hot section temperatures are kept nearly as high as the CFM56 maximum takeoff temperatures. But since the supersonic variant would spend several hours at maximum temperature (compared to jus t a few minutes during takeoff for a subsonic turbofan), this becomes a rather important assumption. Maintaining high temperatures is justified by assuming increased hot section overhaul frequency (not uncommon for a business jet application), and perhaps by offering new turbine airfoils with improved materials , coatings, an d better cooling effectiveness.

Figure 2 . Solid models of the CFM56 - 7B (top) and conceptual modified supersonic variant (botto m ) .

Table 2 . P erformance summary for the STCA derivative engine.

M1.4, 50kft, M0.25, sea Sea level static, ISA level, ISA+27°F ISA+27°F Net thrust, lb /engine 3330 14,140 16,620 Specific fuel consumption, lb /hr/lb 0.943 0.588 0.479 Bypass ratio 2.9 2.9 3.0 Burner temperature, °R 3300 3150 3130 Turbine inlet temperature, °R 3180 3040 3020 Compressor exit temperature, °R 1450 1440 1430 Overall pressure ratio 22 21 21 Fan pressure ratio 2.0 1.9 1.9 Compre ssor pressure ratio 11.2 11.1 11.2 Extraction ratio 1.1 1.1 1.1 Nozzle pressure ratio 5.9 1.9 1.8 Fan pressure ratios are selected such that the nozzle operates on the cusp of choke near sea level. With a small amount of engine derating at low altitude , jet shock cell noise is eliminated during takeoff . A summary of engine pe rformance data is shown in Table 2 . Ambient conditions above 10,000 feet use International Standard Atmosphere (ISA) conditions, while cond itions nearer sea level use hot day (ISA+27°F) conditions. Performance data at sea le vel are shown after engine derating .

Though the supersonic engine variant may appear similar to GE Aviation’s Affinity™ engine [ 33 ] proposed for Aerion Supersonic’s supers onic business jet, it is based on a NASA model predating Aerion Corporation’s public partnership announcement with GE Aviation in 2018. It uses no company - proprietary information. Additional information for the STCA equipped with these engines is available in [ 2 ].

C. Mission performance With aerodynamics, engine thrust and fuel consumption performance known, a mission analysis of the transport can be made using NASA mission performance software [ 34 ]. The design mission is at maximum takeoff gross weight , non - stop, with a single cruise segment at supersonic speed. There are no subsonic cruise segments which might be typical of other missions where supersonic flight might be restricted. Since high - altitude air traffic should be light, block altitude cleara nce is assumed and the airplane is allowed to climb continuously during supersonic cruise. The design mission uses the full payload complement of eight passengers. The mission rules described in [ 35 ] are followed, except that the five p ercent block fuel re serve allowance is omitted, which is more typical of mission rules followed by business jets. The mission profile is shown in Figure 3 . The design mission range of the STCA is 4243nmi.

Performance results for th e STCA using the derivative engine are shown in Table 3 .

4 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference Table 3 : STCA performance.

Takeoff gross wt, klb 121 Operating empty wt, klb 51 Payload, lb 1640 Climb time, min 47 Cruise altitude, kft 44 - 51 Cruise lift - drag ratio 6.6 - 7.9 Block time, hr 5.9 Block fuel, klb 61 Reserve fuel, klb 8 Range, nmi 4243 Figure 3 . STCA mission profile .

III. Noise Prediction Method Another prerequisite for co mputing ANP data for the vehicle is an acoustic model. In general, aircraft system noise prediction tools fall into one of two categories. Both categories are discussed here.

The first category of tools rely on so - called noise - power - distance (NPD) data as the basis for computing noise on the ground . This approach was originally proposed by SAE [ 36 ], revised [ 37 ], and most recently documented by ICAO [ 38 ]. Original equipment manufacturers of civil aircraft supply NPD data for their products so that their noise impact may be modeled. Manufacturers measure NPD data directly, scale or infer them from other NPD data, or predict them from first principles using computational tools in a scenario appropriate for measuring them. System nois e codes based on NPD input data are s aid to be “ICAO Doc. 9911 - compliant” in accordance with [ 38 ]. Examples are the Aviation Environmental Design Tool (AEDT) [ 39 , 40 ] and others [ 41 - 44 ]. These tools, in a sense, start with the answer, because rudimentary noise cha racteristics for each airplane are given as tabular lookup data.

Other noise prediction codes, however, are less empirical and are based more on physics. They are more reliant on first principles, where lossless free - field spe ctra in the vicinity of an air plane are predicted and then propagated in a straight line to a receiver. Examples are NASA’s Aircraft Noise Prediction Program (ANOPP, [ 45 , 46 ] ) and others [ 47 - 50 ]. The distinction between the two catego ries can blur, however, since codes like ANOPP can be coaxed into generating NPD data for use in Doc. 9911 - compliant tools.

All n oise predictions for the STCA ( including its NPD data ) are made using NASA’s ANOPP , with predictive methods selected to represe nt the noise so urces of a supersonic business jet . It should be noted that applying empirical methods that were developed largely for conventional subsonic transports ( powered by high bypass ratio, separate - flow turbofans ) to a supersonic delta - wing airpla ne ( equipped wi th low bypass ratio engine s with supersonic inlet s , long - duct mixer s , and more complex nozzle s ) has a high uncertainty. Sensitivities to uncertainty are assessed in the previous study [ 2 ].

Jet noise is predicted using an empirical method developed by the Society of Automotive Engineers [ 51 ]. For simple, single - stream, round nozzles operating on the cusp of choke, the SAE method is preferred over other available method s based on comparisons of predicted levels to sc ale model and flight test data [ 52 , 53 ]. Since jet noise is typically the dominant source in supersonic applications, it is the subject of recent studies at NASA [ 54 - 57 ].

Broadband fan noise is predicted using an empirical method devel oped by General Electric [ 58 ], and discrete fan interaction tones are predicted using a similar method [ 59 ] (both of these methods are more recent calibrations of ANOPP’s original fan noise me thod). For predicting noise of high - speed, high - pressure - ratio, si ngle - stage fans that might be used in a supersonic application, these methods are preferred based on comparisons made to data collected by NASA: the General Electric High Speed Fan [ 60 ], the Honeywell Quiet High Speed Fan [ 61 ], and a two - stage fan designed a nd built by Pratt & Whitney as part of the NASA - led High Speed Research Program. Fan treatment suppression is estimated using a method developed by General Electric [ 62 ].

Engine core noise is predicted by a method developed by Emmerling [ 63 ]. Engine state da ta computed by NPSS are fed into ANOPP as functions of flight speed, altitude , and engine power setting.

Landing gear, flap, slat and trailing edge airframe sources are predicted using a recalibrated version of the empirical Fink method [ 64 ] . The recalibrat ed model is based on noise measurements of a supersonic delta - wing transport [ 65 ] 5 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference and is documented in [ 66 ]. The predicted levels are adjusted to represent airframe noise levels of a notional supersonic business jet.

With engines mounted above the vehicle, n oise shielding effects must be considered. Shielding (also referred to as barrier attenuation or insertion loss) is an acoustic diffraction phenomenon where sound waves are attenuated when propagated past an impermeable barrier placed between the noise sou rce and an observer. In this study, a simple empirical diffraction model based on optical diffraction theo ry is used. The model was o riginally proposed by Maekawa [ 67 ] and is reproduced in many foundational acoustic textbooks. Shielding is particularly effi cient when the observer is located in the “shadow region” where the noise source is obscured. The delta w ing (see Figure 1 ) provides excellent shielding of forward - radiated fan inlet noise. All other sources are not shielded. Jet noise is a distributed source generated do wnstream throughout the axial exhaust plume. Core noise is predominantly aft - radiating and is assumed to radiate through the exhaust. Fan exit noise also escapes through the nozzle but it is attenuated by treatment in the bypass duct.

Noise levels of all components are predicted as lossless, one - third octave band spectra and are summed in the vicinity of the airplane. The noise sources are predicted at various flight conditions (i.e., the NPD “evaluation points” des cribed in the n ext section ). The noise levels are p ropagated to a pole - mounted receiver located on the ground . NPD data are computed as explained in Section VI . Noise propagation effects include spherical spreading, Doppler shift and convective amplificati on, atmospheric absorption, and ground reflections [ 68 ] based on data for grass - covered ground [ 69 ]. The atmospheric absorption model required for NPD data [ 37 ] differs from the absorption model used for noise certification [ 70 ].

IV. Noise Abatement Departure Procedures ICAO Doc. 9911 - compliant codes require a flight track to be defined in the terminal airspace for every airplane operation so that noise contours on the ground can be computed. Flight tracks may be defined either by fixe d - point profiles or by procedural step calculations using simplified equations of motion and a irplane performance data . Flight tracks for the STCA are defined using the fixed - point profile option. The profile data consist of points defined in space and tim e ( viz. airspeed), and by aircraft configuration and engine power setting. Both departure and arrival profiles are computed .

The state of the engine, the airplane and its flight track have first - order influences on airport - vicinity noise. For Doc. 9911 - com pliant codes to predict operational noise most accurately , the flight tracks should be consistent with those used in real - world operations. Noise Abatement D eparture P ro cedures (NADPs ; sometimes referred to as Noise Abatement Departure Profiles ) are popula r in modern civil aviation, and an airplane such as the STCA would be likely to use them. This section provides an overview of NADPs for the STCA.

Operational NADPs differ from takeoff procedures used in noise certification. R eference procedures of Annex 16 noise regulations [ 12 ] are generally more restrictive than procedures ordinarily found in operational practice. For example, noise certification reference procedures described in Section 3.6.2(d) of the Annex require a const ant climbout speed, wh ereas NADPs typically include at least one in - air acceleration segment. And i n practice, p ilot - initiated engine power cutbacks differ from those permitted in Section 3.6.2(b) . In noise certification , power cutbacks are usually much de eper .

There are other differences as well, including two that might be unique to supersonic transports (though at this writing, applicability to subsonic transports is under discussion) . The first pertains to engine thrust management.

Section 3.6.2(a) of the Ann ex requires that takeoff thrust be maintained from brake release until the point where the pilot - initiated thrust cutback is permitted. However, a programmed lapse rate (PLR) procedure is under discussion among regulators for future supersonic aircr aft [ 15 ] . The PLR procedure is a thrust lapse programmed to occur automatically at low altitudes. It is not to be confused with the pilot - initiated engine power cutback occ urring later. It would be employed both during noise ce rtification and in normal operations. In certification, a PLR procedure would reduce noise at the lateral certification monitor, and in operation it would reduce noise near airports . A more detailed di scussion of PLR procedures and how one is envisioned to be implemented for the STCA can be found in Ref s . [ 2 ] and [ 19 ] .

The second difference pertains to airframe configuration changes. Section 3.6.2(e) of the Annex requires the position of high - lift de vices be maintained throughout the reference procedure . This differs from operational practice, where flaps and slats are retracted on sensible schedules. I n a manner similar to the PLR procedure, i t should be possible to program flaps to retract automatic ally .

Both of these procedures could be implemented as a Variable Noise Reduction System ( VNRS ) under the provisions of [ 71 ]. Normally, r egulating authorities might be reluctant to approve any pilot - initiated procedure that would increase the workload of t he flight crew at low altitudes . But i t is thought that exceptions w ould be permitted 6 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference if computer - controlled automatic throttle scheduling and flap retraction are used, making pilot initiation unnecessary.

An automatic digital engine control implementation of these procedures could use an airplane’s weight - on - wheels sensors, a irspeed, altimeter, attitude and air temperature indicators, or perhaps airport navigational aids to begin preprogrammed actions. A companion paper describing the use of a VNRS to mini mize noise of the STCA in a certification scenario has been written [ 19 ].

The methods and tools used to derive departure profile data for the STCA are selected by NASA. Derivation of data for other aircraft are likely to use ot her methods, tools, and selection criteria. This paper documents NASA’s approach .

A. NADP Characteristics ICAO [ 72 ] and FAA [ 73 ] recommend NADPs for operators of turbine - engine airplanes over 75,000lb. Procedures must be consistent with existing airworthiness s tandards. No more than two procedures for each aircraft type are recommended.

Profiles having a close - in noise benefit are intended to abate noise for communities close to airports. These profiles are often char acterized by an e arly pilot - initiated engine power cutback and an initial c limb at constant airspeed with takeoff flaps deployed . From that point, t he profiles continue typically with a flap retraction and an acceleration segment. These procedures are named NADP - 1, c lose - in NADP, or ICAO - A . T he latte r name is used in AEDT , though the terminology is now obsolete.

Alternately, p rofiles having a distan t noise benefit are intended to abate noise for communities far from a irports .

These profiles are usually characterized by an e arly flap retraction and an acceleration segment . These procedures are named NADP - 2, d istant NADP, or ICAO - B (now obsolete).

In general, close - in procedures tend to devote more of the vehicle’s total available energy towards increasing its potential energy (i.e., altitude) than dista nt procedures do. A generalized schematic of noise abatement departure pr ocedures is shown in Figure 4 .

Note that the PLR procedure is indicated on the figure. A discussion of the PLR procedure as implemented by NASA for the STCA is can be found in [ 2 ]. The STCA is assumed to have completed its PLR procedure before specific NADP steps begin. The goal of this study is to select one close - in benefit pro file (in the style of NADP - 1) and one distant benefit profile (in the style of NADP - 2) for the STCA .

Figure 4 . NADP schematics.

7 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference B. Survey of the Aircraft Noise and Performance Database Most of the details of NADPs are unspecified and are left to manufacturers to determine. A ccelerati ons, changes in configuration, engine power and the sequence and timing of events are not always consistent from manufacturer to manufacturer, or even from product to product. With so many v ariables involved, it is not immediately obvious how the STCA shou ld be flown. The refore, a survey of the Aircraft Noise and Performance database (defined in a Doc.

9911 appendix and available online [ 74 ]) is used to guide the selection of profile paramete rs for the STCA ’s operational profiles. Nomenclature in this survey correspond s to that used in AEDT. Departure procedures considered are ICAO - A, ICAO - B, and DEFAULT (also known as STANDARD). Some aircraft types are omitted from the survey.

Each d eparture procedure in the ANP database is examined for its cutback trigger ( i.e., the criterion that determines when engine power is reduced), cutback level (the reduced engine power setting), flap retraction trigger (the criterion for flap retraction), acceleratio n trigger (the criterion for one or more acceleration segments ), an d airspeed s . Note that a ny ICAO - A, ICAO - B, or DEFAULT procedure in the ANP database can be classified as either a close - in noise benefit or a distant noise benefit. From the survey, the following observations are made: 1. ICAO - A and NADP - 1 procedures are ide ntical.

2. ICAO - B and NADP - 2 procedures are nearly identical, except that : o ICAO - B cutback s occur after the acceleration segment.

o NADP - 2 cutback s occur “during the flap/slat retraction seque nce at a point that ensures satisfactory acceleration performance” [ 72 ].

3. DEFAULT procedures can be classified as either close - in or distant (but nearly all of them a re distant).

4. In modern operational practice, “distant benefit” departures are commonly flown as NADP - 2 [ 75 ].

5. Aérospatiale/BAC Concord e exits the terminal airspace (i.e., 10,000ft above field elevation) at 2 50 k c as.

6. 10 airplane types exit the terminal airspa ce at speeds above 250kcas.

7. One 747 type exits the terminal airspace at 300kcas.

The latter three items are relevant for supersonic transports , which typically have a preference for high airspeed at low altitudes. These observations are important in setti ng NADPs for the STCA, as discussed in the following section.

C. Parametric Variation and Selection of Departure Variables A flowchart d escribing how the STCA may be flown using different procedures is adapted from [ 75 ] and is shown in Figure 5 . Major independent variables are parameterized; these are labeled using the letters “ A ” throu gh “ N ” in the flowchart. Rather than allow the independent variables to be continuous - real , they ar e assigned discrete values with the intent to perform a n optimization based on a factorial design of experiments . This type of grid - based optimization is an “exhaustive search.” The cost of this brute - force strategy is higher than a search aided by an opti mizer. But if each objective sample runs reasonably fast, it is attractive since finding the global optimum is virtually ensured (provided the grid intervals are sufficiently fine). The optimal NADP design s are selected simply by inspection. The danger of an optimizer becoming “stuck” on a local optimum is thus avoided. The left side of the flowchart is intended to represent departure procedures patterned afte r NADP - 1/ICAO - A; while the right side represents procedures are patterned after NADP - 2/ICAO - B. The independent variables are permuted using discrete values shown in the figure, and each resulting NADP profile is examined for merit as explained below.

For the STCA, a ll NADP procedures begin at 400ft above field elevation. This is at the beginning of the second climb segment when the airplane has reached its takeoff safety speed plus 35kt (i.e., V +35kt , as explained in [ 2 ] ). By that altitude, it has completed its programmed thrust lapse and its gear has been fully retracted. All takeoff profiles are evaluated at maximum gross weight from a sea level field.

Three pilot - initiated engine power cutback level s are investigated (see Figure 5 ) . The range is defined by the thrusts required to maintain climb gradient s of 1.5, 2.0, and 2.5% with one engine inoperative ( note that 1.5% is the minimum clim b gradient from 400ft to 1500ft permitted for trijets per §25.11 1 of [ 76 ] ).

Two terminal airspeeds are investigated ( these are airspeeds achieved upon exiting the terminal airspace at 10,000ft above field elevation) . Referring to Figure 5 , t he exit airspeeds for the STCA are either 250kcas or 29 0kcas. 250kcas is a logical choice because, unless otherwise authorized, it is the maximum airspeed allowed under 10,000ft in the U.S. per §91.117 of [ 77 ] . Other countries have also adopted this speed limit. Thus, exiting the terminal airspace at the 290kca s higher speed requires some justification.

8 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference Figure 5 . Flowchart formulation and parametrization of variables.

H igher exit speeds are found in the ANP database survey , as discussed in the section above. And the STCA, like many supersonic airplanes, has a preference for high airspeeds at low altitudes. A chart showing the STCA’s thrust demand is shown in Figure 6 for level and steady flight at maximum weight at an altitude of 2000 feet. Noted in the figure is the takeoff safety speed (V ), determined by a one - engine inoperative, balanced field calculat ion following the guidance in [ 76 ], and the 250kcas speed limit.

The STCA’s minimum drag speed is nearly 280kcas. Speeds below this are in the so - called region of r eversed command, where to fly more slowly requires more thrust to overcome Figure 6 . Thrust demand of the STCA in level, steady flight .

increasing lift - dependent drag. This is in contrast to a similarly - sized subsonic airplane with takeoff flap deflections in these conditions, which may have a minimum drag speed of perhaps only 200kcas. Flying safely in this region requires adequate thrust margins, shown in the figure as the diffe rence between available thrust and required thrust with all engines operating and with one engine inoperative. Supersonic transports are li kely to have thin wings with low aspect ratios and simple flap systems. They are likely to require high takeoff speed s before sufficient lift is generated to lift off, and even higher airspeeds to climb with significant thrust margin. In contrast with most subsonic transports, in - air acceleration segments to high airspeeds are preferred for supersonic transports. In oper ational practice, it was common for the Concorde to accelerate and to reach 250kcas by 400ft above field elevation [ 74 ], and similar in - air accelerations were planned for the proposed supersonic High - Speed Civil Transport [ 78 , 79 ] .

In this study, 290kcas is justified since it is close to the STCA’s min imum drag speed.

9 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference With the parameterization of departure variables chosen (see Figure 5 ) , 48 and 216 possible combinations are identified for NADP - 1 and NADP - 2 profiles, respectively. Each of these profiles is analytically “flown” using a forward - differencing trajectory solver th at satisfies the equations of motion at discrete intervals [ 80 ]. Using the noise methods described in Section III, the sound exposure l evel (SEL) noise metric is evaluated for ground observers located directly under the flight path of each trajectory. A ne w noise metric is created to determine the merit of eac h profile . The metric consists of the area under the curve of SEL and distance f rom the airport (in units of dBA - nmi).

The curve is weighted to emphasize SELs close to the airport, or far from the airport, depending on whether NADP - 1 or NADP - 2 profiles are assessed. Additional preference is given to profiles exiting the terminal airsp ace at 290kcas.

The best profiles are chosen by inspection. The preferred NADP - 1 and NADP - 2 profiles for the STCA ar e shown in Figure 7 . F or comparison purposes , t he Annex 16 noise certification profile computed in [ 2 ] is also shown. Also shown in the figure are “evaluation points” for computing noise - power - distance (NPD) data. They are roughly representative of the flight conditions for any departure procedure. These NPD evaluation points will be discussed further in Section VI.

V. Influence of Stage Length on NADPs The NADPs determined for the STCA in the previous section are evaluated at maximum takeoff gross weight. But of course in operational practice, airplanes do not always depart at maximum gross weight. And s ince gross weight influences an airplane’s takeoff trajectory, additional NADP profiles at gross weights less than maximum are required.

The l ogic to determine these weights and the takeoff profiles associated with them are discussed in this section.

When making noise evaluations, AEDT models operations be tween airport pairs according to trip distance intervals. For simplicity, o perations are ca st into discrete bins called stage lengths. When a code such as AEDT is used to perform a multiple - event fleet noise analysis, a stage length must be assigned to eac h operation. Stage lengths consist of representative trip distances and the corresponding t akeoff gross weights needed to fly those distances. Stage length is a surrogate indicator of takeoff gross weight. Using the same departure procedures defined in the previous section , additional profiles must be developed for the stage length missions at l ower gross weights . Stage lengths defined for the STCA are shown in Table 4 . A takeoff gross weight is sought for each stage length so that additional NADPs can be computed for them.

To determine the weight and fuel required for ea ch stage lengt h , an optimized mission planning calculation is required . Mission range performance is a strong function of fuel load, takeoff gross weight, and how the mission is flown. Cruise variables for the STCA are selected that optimize each mission. Logically, t he optimization variable for a supersonic business jet like the STCA is block time .

However, depending on the distance flown, additional considerations are given for block fuel, climb time, and practicality.

The following mission constraints ar e assumed. Breguet climbing cruise i s assumed at altitudes over 40,000ft, where block altitude clearance is expected. The crew is assumed to use on - board programmed guidance to manually achieve a climbing cruise, or perhaps it will be automatically program med into airplane flight management computers.

Figure 7 . Departure profiles for the STCA .

The maximum subsonic and supersonic Mach numbers 10 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference allowed are 0.92 and 1.4, respectively. Fixed altitude cruise is assumed under 45,000ft. Supersonic speed is allowed only over 40,000ft. Cruise altitude and spe ed are optimized for each stage leng th, using a composite objective consisting of block time and block fuel. Results are discussed below.

1. Primary Stage Length Performance: S upersonic airplane s like the STCA (with no features to reduce its sonic boom) are likely to have restrictions on supe rsonic overland flight. A realistic mission between a specific city pair might consist of multiple cruise segments and speeds to avoid flying at supersonic speeds where prohibited . Indeed, route planning for supersonic a irplanes can become quite complicate d [ 11 , 81 ]. But to evaluate noise using a tool like AEDT, the stage lengths defined for each vehicle type are generic (i.e., there are no specific city pairs associ ated with any stage distance ). For a subsonic airplane, this is not usually a concern. But for the STCA , stage length performance differs significantly when restrictions are placed on cruise speed. In this study, the modeling shortcomings are simply accept ed. Mission assessments are made strictly on the basis of overall vehicle performance .

With no such restrictions, s upersonic Breguet cruises at Mach 1.4 are preferred for the three longest stage lengths (2200, 3200, and 4240nmi) . However, as trip distance becomes shorter, the block time metric shows diminishing returns and it becomes sensible to consider block fuel as an additional measure of merit. A s ubsonic cruise at Mach 0.92 is preferred for the three shortest stage lengths (350, 850, and 1350nmi) . Since a supersonic business je t operator would presumably fly most often at supersonic speeds when appropriate, this stage performance is named primary stage performance. These are shown on the left side of Figure 8 .

2. Alternate Subsonic Stage Length Performance : Unlike subsonic airplanes, supersonic airplanes have multi - Mach cruise speed options. Even when flying over water with no restriction on speed, the operator of a supersonic airplane has the option of flying at subsonic speed. In the case of t he STCA , it s subsonic cruise range is substantially greater than its supersonic cruise range (note this may not be true of all supersonic aircraft). B ut g iven th e option , it is possible that an operator might elect to save fuel or to fly further by cruising at a subs onic speed at the expense of time . Thus, alternate ranges are given for the STCA where all - subsonic missions are flown for the three long er stage lengths. A bove 40,000ft , block altitude clearance is assumed and a Breguet cruise is allowed . Alternate subson ic stage length performance is shown on the right side of Figure 8 .

P erformance for all stage lengths is shown i n Table 4 .

3. Departure Engine Power Management At each gross weight, engine pow er management must be defined. Beginning with AEDT version 3b in 2019 , the option of reduced thrust takeoffs became available (See Section 3.6.2.3 of the AEDT3b Technical Manual [ 39 ] ). The reduced thrust setting models part - pow er takeoffs ( that are common in practice ) when maximum thrust is not required.

However, this option was not used in the CAEP supersonic exploratory s tudy. For the CAEP study , maximum takeoff thrust available wa s used from the beginning of the takeoff roll to the pilot - initiated cutback. Field lengths and climb performance were allowed to vary.

Table 4 . STCA preferred stage lengths.

11 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference Figure 8 . STCA primary stage length performance (left); alternate subson ic stage length performance (right).

VI. Noise - Power - Distance Data ANOPP is used to predict all four types of NPD data used by AEDT: sound exposure level, maximum A - w eighted level, maximum tone - corrected perceived noise level (PNLTmax), and effective perceived noise level (EPNL). And s ince ANOPP operates on the one - third octave band paradigm, all spectral data (computed natively by ANO PP) are also available for use in AEDT. The airplane is flown over an observer at altitudes ranging from 200ft to 25,000ft at te n discrete engine power settings. The simulated pole - mounted receiver is located 4ft above ground level. Ground reflection calculations are enabled. The atmospheric absorption model [ 37 ] required by AEDT is applied to the propa gation calculations.

N oise sources are functions of airspeed and altitude . Jet noise, for example, is caused largely by turbulent mixing of the jet wit h ambient air . Strength of the s hear layer is reduced as the airplane flies faster, and so jet noise is s trongly influenced by airspeed. And naturally, airframe noise varies strongly with airspeed. Indeed, all sources are influenced by convective amplifica tion and Doppler effects that vary with airspeed. So , in order to predict noise accurately for NADP profi les, noise predictions should be made at airspeeds that are most applicable to flight conditions of the NADPs. Thus, with the noise prediction methods described in Section III, and with the flight conditions determined for NADPs in Section IV, noise - power - distance data for the STCA can be calculated.

If NPD data are computed via ANOPP, appropriate airspeeds must be provided by the user for each value of corrected net thrust per engine . Since thrust is one of the independent variables of NPD data, t he airspe eds should be values that are most appropriate for the thrust in question. For the STCA, these airspeeds are at the profile conditions labeled “evaluat ion points” in Figure 7 . The NPD evaluation points are selected to be roughly re presentative of the flight conditions encountered during any departure procedure.

But t h e reference airspeed for NPD data in all Doc. 9911 - compliant tools is 160 ktas . In other words, whenever a compliant tool such as AEDT queries an NPD table, the noise me tric returned is expected to be appropriate for a flight profile segment “flown” past a receiver at 160ktas. If the flight segment of the user - specified profile is at some other airspeed (and it usually is) , AEDT makes an adjustment to the duration - depende nt metrics SEL and EPNL. Therefore , before SEL and EPNL values computed by ANOPP can be ready for use in a code like AEDT, a “counter adjustment ” must be made to the m . This correction is a post - processing a djustment is added to the SEL and EPNL levels afte r they are computed by ANOPP (the adjustment is 10log [VTAS/160ktas] , where VTAS is the true airspeed of the NPD evaluation point ).

Addi ng more independent variables can increase the accuracy of NPD data . Multi - speed NPD data , for example, would include a irspeed as an additional independent table lookup variable. This would ensure more accurate modeling of noise source s that are strong functions of airspeed, such as j et noise and airframe noise. The STCA, for example, 12 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference might benefit from additional NPD data above 250ktas (see Figure 7 ). Research is ongoing [ 75 , 82 ] that may result in addi ng airspeed - dependent NPD data to future versi ons of AEDT.

Corrected net thrust levels are chosen to cover the range of engine operation expected in practice. The NPD data for the STCA are shown in Figure 9 . They are ready for use in AEDT. The data were provided to CAEP analysts performing the supersonic exploratory study.

The STCA’s NPD data (unadj usted for airspeed) can be used to check the Annex 16 certification noise levels made in a previous ANOPP simulation ( see [ 2 ]). Table lookup errors for EPNL and PNLTmax for flyover and approach conditions are less than 1dB. The se lookup errors are attributed to the following modelin g differences between the NPD simulation and the Annex 16 certification noise simulation: • Layered atmosphere at ISA+18 ° F (certification predictions) vs. a homogeneous atmosphere at ISA (NPD prediction s) .

• SAE ARP 866A absorption (certification predictions) vs. SAE AIR 1845A absorption (NPD predictions) .

• More realistic profiles with variable engine thrust (certification predictions) vs. level, steady flyovers at constant engine thrust (NPD predictions) .

Figure 9 . STCA n oise - power - distance data.

VII. Summary A notional 55 - tonne supersonic technology concept airplane is discussed in this paper. It has been developed and studied by NASA at the request of the FAA and ICAO. The behavior of the STCA in an operational setting is considered. A ircraft noise and performance data consisting of n oise abatement departure procedures , stage length performance, and noise - power - distance data are predicted . NASA tools are used to compute all data . Limitations are discus sed in the paper. Also investigated are som e of the ant icipated behaviors and requirements of supersonic aircraft in the terminal airspace. The influence that future supersonic airplanes will have on operations , fuel consumption, airport noise, and air quality has been assessed by a CAEP supersonic exploratory study . T he STCA serv ed as an analytical proxy for similar early market entrant supersonic airplane types. Being nonproprietary and transparent, the STCA is ideal for use in ICAO’s public studies. The data presented in this paper were provided to CAEP anal ysts for the study.

13 of 16 American Inst itute of Aeronautics and Astronautics 28th AIAA/CEAS Aeroacoustics Conference VIII. Acknowledgments Thanks to NASA’s Commercial Supersonic Technology Projec t for supporting this study. Thanks also go to the International Coordinating Council of Aerospace Industries Associations , particularly GE Aviation an d Gulfstream Aerospace Corporation, for their helpful guidance and suggestions. Thanks go to ICAO’s Working Group 1 (Aircraft Noise, Technical) , and to the FAA, for including supersonic research into their work programs. Together, they are helping the worl d understand the impacts of future civil supersonic transports. The vision of viable supersonic transports operating responsibly in an environmentally sustainable manner is a shared pursuit.

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