Design of a Short-To-Medium-Range Baseline Aircraft with an Entry into Service in 2035 for the HOPE Project

Abstract

When assessing new technologies at the overall aircraft level, it is crucial to establish an appropriate benchmark to evaluate the resulting performance of the final concept aircraft. This publication defines a short-to medium-range baseline aircraft for entry into service in 2035, which serves as a benchmark for the HOPE project. This aircraft is sized using the Bauhaus Luftfahrt Aircraft Design Environment (BLADE), and its engine model is derived using the in-house tool Aircraft Propulsion System Simulation (APSS). The top-level aircraft requirements and technology assumptions for the entire project timeline are also defined and used for the baseline aircraft. Compared to a state-of-the-art aircraft (entry into service in 2015), the baseline aircraft consumes 21.4% less block fuel during the design mission and 20.9% less block fuel during the typical mission, since its operating empty mass is reduced by 13.6%, its aerodynamic performance is improved by 8.5%, and it has 6.4% more efficient engines.

1. Introduction

In light of the Paris Agreement, the aviation sector has set ambitious goals to reduce its climate impact, targeting net-zero emissions by 2050 [1]. Recent analyses have shown that aviation caused 3.5% of the effective radiative forcing between 1940 and 2018 [2]. If no mitigating actions are taken, this number can be expected to increase since the number of passenger journeys is forecasted to double by 2040 [3], caused by a forecasted growth of 8.4% p.a. until 2027 and 3.6% p.a. from 2027 until 2043 [4]. One major pillar of the endeavor to reduce the climate impact of aviation is the development and integration of new technologies, enabling more energy-efficient aircraft alongside reducing the climate footprint through the use of alternative fuels.

In this context, hydrogen has regained attention as an energy carrier for aviation. However, its integration into aircraft and operations remains challenging. The Hydrogen Optimized Multi-Fuel Propulsion System for Clean and Silent Aircraft (HOPE) project, funded by the Horizon Europe program of the European Union, aims to bridge the gap until the comprehensive introduction of hydrogen into the aviation sector through the development of a dual-fuel aircraft. This aircraft features liquid hydrogen and sustainable aviation fuel-burning dual-fuel turbofans and a boundary-layer-ingesting (BLI) fan driven by a hydrogen fuel-cell auxiliary propulsion and power unit, with an entry into service in 2035 [5].

To evaluate the HOPE aircraft with the aforementioned technologies, the definition of a conventional baseline aircraft that is sized to the same top-level aircraft requirements (TLARs) and technology assumptions is necessary. This paper describes the specifications and performance of such a baseline aircraft. The methodology used to derive the baseline aircraft and TLARs is stated, followed by a description of the overall aircraft geometry, mass breakdown, aerodynamic and engine data, and low- and high-speed performance. To demonstrate the design sensitivities of the baseline aircraft, selected linear trade factors are presented in the conclusion of this paper.

2. Methodology

This study is conducted by creating a derivative of a known aircraft, which is called the HOPE reference aircraft (HOPE-R) in the following chapters. The HOPE-R is a project-internal Airbus-A320neo-class aircraft, which is based on and calibrated to publicly available geometric, performance, and mass data [6]. This aircraft is then modified using technology assumptions alongside the integration of technology bricks for an entry-into-service (EIS) in 2035 to derive the HOPE baseline aircraft, which is designated as HOPE-BL. Therefore, the HOPE-BL features a low-wing, conventional tail, tube-and-wing configuration, with its engines and main landing gear mounted to the wings, same as the HOPE-R. As the goal of the project is to evaluate the dual-fuel propulsion system alongside BLI propulsion, configurative changes to the aircraft are minimized to ensure comparability.

To explain the methodology behind the derivation of the HOPE-BL, the Bauhaus Luftfahrt Aircraft Design Environment (BLADE), which was used to conduct all studies, is introduced. Afterward, the engine design framework is summarized. Finally, the TLARs and technology assumptions for the HOPE-BL are presented to define the starting point for subsequent studies.

2.1. Bauhaus Luftfahrt Aircraft Design Environment

BLADE is a modular in-house aircraft design software that has been specifically developed to feature a traceable and automated aircraft design process while applying modern software design principles, including extensive testing. Within BLADE, Python modules dedicated either to an aircraft design discipline or an aircraft component exchange data via the Common Parametric Aircraft Configuration Schema [7] and are sequentially executed during an iterative design process. The underlying methods and heuristics of the BLADE modules, as well as the overarching methodology used in this paper, are explained in the following paragraphs as used for the design of the specific aircraft configuration for this study.

The overall aircraft design methodology employed in this study is based on two fundamental procedures: calibration and sizing. First, the HOPE-R is modeled in BLADE, such that the employed methods are calibrated based on published geometric, mass, performance, and aerodynamic data, and then refined based on internal estimations. The data used in the calibration can be found, for instance, in the Aircraft Characteristics, Airport and Maintenance Planning (ACAMP, [6]) document and the CSMR-01 aircraft model [8]. Furthermore, the payload-range diagram found in the ACAMP document is utilized to achieve a match between the published aircraft performance and its corresponding modeling. Therefore, the airframe is sized to match a given operating empty mass (OEM). Subsequently, the thrust-specific fuel consumption (TSFC) of the engine is iterated to meet the desired maximum take-off mass (MTOM) while keeping the aerodynamics fixed. These calibrations are applied using linear scale factors, either on a single parameter, e.g., scaling the wing mass model output to match the wing mass of the real aircraft, or sequentially on a set of dependent parameters, e.g., first scaling the zero-lift drag coefficient of each component to achieve a desired drag breakdown and then scaling the overall aircraft drag coefficient to match certain aerodynamic characteristics. After the reference aircraft has been modeled and calibrated, a set of technological improvements is implemented, as discussed in Section 2.5, and the baseline aircraft is sized based on the requirements defined in Section 2.4 and Section 2.6. This aircraft design methodology is illustrated in Figure 1, and the sizing procedure is analyzed further in Section 2.6.

In the following subchapters, the dedicated BLADE modules for different aircraft components, namely the fuselage, the wing, the empennage, the landing gear, and the engine, alongside its nacelle and pylon, are described. The introduction of discipline-based BLADE modules follows afterward.

2.1.1. Fuselage Module

The front and aft fuselage sections are modeled using discretized elements to represent a predetermined shape approximating the real aircraft nose and tail cone, whereas the constant section is modeled using a cylinder to fit a defined cabin length.

The structural mass of the fuselage is calculated using the semi-empirical equation of Part C stated in [9]. The x-position (longitudinal) of the center of gravity (CG) of the fuselage is calculated according to Chapter 7 of [10], and the z-position (vertical) is determined by an in-house-derived factor of the fuselage height.

2.1.2. Wing Module

The double-trapezoid wing geometry is based on the trapezoidal wing definition found in Chapter A-3.1 of [11], which is also used to define the wing reference area. The wing is scaled using a defined wing loading (see Section 2.6) and positioned with a specified static margin calculated using Equation E-50 presented in [11].

The semi-empirical method described in [12] is used to estimate the wing structural mass. The logic of Tables 8–15, as presented in [11], which includes approximate locations for the CG location of each component, is used for the determination of the CG x-position, and the z-position is estimated with an in-house formula, which is based on the wing span, y-position, z-position of the wing root, and wing dihedral. The mass of the winglet plus the additional structural wing mass is calculated following the procedure presented in [13]. The CG x- and z-positions of the winglet are estimated based on its geometrical center.

2.1.3. Empennage Module

The tail surfaces are geometrically defined by single trapezoids using the same method as for the wing. Both surfaces are sized and positioned by keeping their tail volume coefficients and the 25% mean aerodynamic chord (MAC) locations constant. Hence, the vertical tail is scaled using Equation 6.26, and the horizontal tail is scaled using Equation 6.27, as found in [14].

The horizontal tailplane (HTP) mass is estimated using the semi-empirical formula in Part D of [9]. The CG x-position is determined in the same manner as that of the wing, and the CG z-position is predicted using an in-house method that takes the HTP root position, span, and dihedral into account. The mass of the vertical tailplane (VTP) is calculated using the semi-empirical formula in Part E of [9]. The estimation of its CG x-position is analogous to that of the HTP. The z-position is determined using an in-house method based on the root position and span.

2.1.4. Landing Gear Module

Regarding the geometry of the nose or main landing gear, a single cylinder reflects its pistons and struts, and axles and tires are represented by cylinders. The dimensions of the wheels are estimated using the landing gear load calculation methodology defined in [14]. The strut and axle radii are scaled with the wheel diameter, and the axle length is scaled with the wheel width. Furthermore, the rake angle can be set for the nose and main landing gears. The length of the landing gear is determined using a requirements-based approach based on [15], which evaluates the requirements for longitudinal stability, lateral stability, and tail strike angle. Additionally, if necessary, the main landing gear x-position is adapted to ensure a longitudinally and laterally stable aircraft. The positions of the fuel tanks, leading/trailing edge devices, and available support/auxiliary spar placement are considered when positioning the main landing gear. Optionally, the rake angle can also be altered to meet the longitudinal stability requirement.

After the geometry and positions of the landing gear are determined, the masses can be estimated. The mass of the nose landing gear is estimated using the semi-empirical formula stated in [16], whereas the mass of the main landing gear is estimated using the semi-empirical formula presented in [17]. The CGs are derived from the individual geometric centers of the cylinders.

2.1.5. Engine, Nacelle, and Pylon Module

The nacelle is scaled in diameter following the semi-empirical procedures presented in [18], as well as in length according to the fan diameter, based on [19], and positioned relative to the wing by specifying a spanwise position and an offset in the x- and z-directions to the corresponding wing leading-edge position.

Both the nacelle and engine CGs are specified as inputs. In the studies conducted in this work, the CG for both components was set to 40% of the engine length relative to the inlet, representing a typical value for turbofan engines (cf. [11,19,20]). The masses of the nacelle, dry engine, and engine subsystems are calculated using the semi-empirical methods described in [19]. To accommodate the mass changes encountered during the aircraft design process, the mass is scaled based on the engine fan diameter and standard day-corrected core engine mass flow. The associated engine performance prediction is based on pre-calculated engine performance decks (see Section 2.2), which are scaled during the aircraft design process to a given design thrust, i.e., the top-of-climb (TOC).

Trapezoidal wings are used to reflect the geometry of the attached pylon, which is positioned relative to the nacelle and wing. The semi-empirical method described in [9] determines the mass of this component, and the CG equals the geometrical center of the pylon.

2.1.6. Overall Aircraft and Subsystem Mass Estimation Module

Aircraft subsystems are currently not modeled geometrically within BLADE; therefore, the mass and CG estimation is included in a module, which also calculates the design masses and CGs of the overall aircraft by summing the component masses. This module is presented in the current subchapter, transitioning the explanation from the component-based modules to the discipline-based modules of BLADE. Regarding its overall functionality, the module is also used to synthesize individual fuel masses and CGs from aircraft components.

The corresponding subsystem masses are estimated using the semi-empirical equations presented in [21]. The CGs are based on in-house methods. The x-positions of the air conditioning, de-icing, flight controls, hydraulic, and electrical systems are determined using wing geometry parameters, such as the root position, root chord, span, dihedral, and length and position of the MAC. The z-positions of the air conditioning and electrical systems are estimated using the fuselage height, whereas the de-icing system, flight control, and hydraulic systems are dependent on the wing geometry parameters, as explained above. The auxiliary power unit, avionics, and instrument panel are purely dependent on fuselage geometry parameters, such as overall length, tail length, and height.

The maximum landing mass (MLM) is calculated according to Chapter 8.2.4 of [11], and the resulting CG is derived using the reserve fuel mass estimated with the mission analysis module, zero fuel mass, and their CGs. The maximum fuel mass and CG are derived from the associated fuel tank masses and CGs. The maximum zero fuel mass and CG are obtained from the OEM and maximum payload mass, as well as their CGs.

2.1.7. Overall Aircraft Aerodynamic Analysis Module

The prediction of the aerodynamic performance of the overall aircraft is divided into low- and high-speed drag polars. The drag is predicted by estimating the induced, profile, interference, wave, leakage, and protuberance drag. The profile drag can be further divided into lift-dependent and lift-independent profile drag fractions.

The high-speed polars are calculated for a target aircraft lift coefficient, ISA deviation, altitude, and Mach number. The low-speed polars are calculated using the methodology of the high-speed polar calculation and adding increments for the used aircraft configuration regarding the high-lift devices and landing gear. Therefore, in the following paragraphs, first the calculation of the high-speed and then the calculation of the low-speed polars are explained. As these associated methods are complex, readers should refer to the cited references for a complete and detailed explanation and discussion.

With the estimated wing lift curve slope according to Chapter 12 of [14], the fuselage-corrected aerodynamic center can be estimated according to Appendix E of [11]. Using this information, the untrimmed wing-fuselage pitching moment of the same reference alongside the user-defined reference CG can be calculated, based on the design cruise condition with a relative fuel loading of 53% for this paper. This enables the calculation of the moment balance between the HTP and wing, yielding the corresponding trimmed lift coefficients for each lifting surface. Similar to the wing lift curve slope calculation, the induced drag is estimated according to the proposed procedure by [14] described in Chapter 12, which is based on the subsonic drag predictions by [11] in accordance with Appendix F and the supersonic calculations of Chapter 12 of [14]. The component-based profile drag estimation follows the methods described in Appendix F of [11] using the friction drag coefficient of [14] (Chapter 12). These drag components are then corrected to reflect the interference drag by [14], as stated in Chapter 12. The wave drag is predicted in accordance with Lock’s approximation [22,23] above the critical Mach number, where the drag divergence Mach number is estimated using [24,25]. Below the critical Mach number, wave drag is set to zero by assumption. Finally, the leakage and protuberance drag is calculated according to Chapter 12 of [14]. The trim drag, which is considered intrinsically during the polar calculation, is determined separately in a post-processing routine.

The low-speed maximum lift coefficient is calculated according to Appendix E of [11] for the basic contribution and Appendix G is used for the flap- and slat-induced increments, whereas the drag increment prediction follows Appendix G of [11] using the friction drag coefficient calculation of [14], as explained previously.

The associated aircraft angle of attack is calculated using the target lift coefficient, the lift coefficient for zero angle of attack, and the lift curve slope of the aircraft.

2.1.8. Mission Analysis Module

Finally, the synthesis of all module data enables the simulation of the designed aircraft flying specified still-air missions complying with regulations [26]. This mission analysis estimates the fuel consumption for a given mission range and payload by creating a mission trajectory that is discretized into mission points and solving the equations of motion sequentially for each mission point [27,28]. Each mission segment can be discretized based on time, altitude, or aircraft mass, such that the most appropriate discretization can be selected to achieve a balance between computational efficiency, stability, and accuracy (e.g., climb is discretized with altitude, whereas cruise is discretized with aircraft mass). At all mission points, the aircraft is represented by a point mass, for which the force balance is solved to determine the acceleration, energy demand, and flight path angle. For segments such as take-off and climb, a thrust rating is utilized, and thus, the acceleration and flight path angle are calculated based on the resulting force balance. For cruise flight points, a steady-level flight is assumed, and the required thrust is calculated based on the aerodynamics of the aircraft. For descent, a constant flight path angle is enforced, and an idle rating is applied to determine the resulting acceleration and fuel consumption.

Additionally, the field performance is predicted by sequentially solving the mission points belonging to the take-off phase and running a take-off speed optimization algorithm, which determines the decision, rotation, and take-off safety speeds to minimize the take-off field length for a given aircraft configuration. This optimization aligns with the procedures described in [29]. Furthermore, during the overall mission analysis, an initial cruise altitude optimization and a step-cruise procedure, which are based on specific-air-range evaluations at different altitudes, are applied to minimize fuel consumption and simultaneously reflect real-world operations.

2.1.9. Aircraft Design Workflow

All previous modules are finally connected to each other and executed sequentially in a loop until convergence is reached. It is possible to define iteration parameters that are varied automatically to meet a specific value (e.g., automatically changing the wing area to reach a specified approach speed). The definition of multiple target parameters is possible; in this case, all targets are evaluated at the end of each iteration, and the most stringent criterion is used to determine the input value guesses for the next iteration until all target parameters satisfy the required values. Therefore, the creation of carpet plots or full-factorial studies is not required when conducting aircraft design studies using BLADE. The specific set of requirements and their target values for this study are presented in Section 2.6.

2.2. Engine Design Framework

Propulsion system modeling is performed using the Bauhaus Luftfahrt in-house tool, Aircraft Propulsion System Simulation (APSS) [30,31,32]. APSS can size conventional and novel propulsion systems (design) and calculate their respective operational performance (off-design). The software is fully developed in MATLAB^® and uses thermodynamic property tables based on NASA’s Chemical Equilibrium with Applications (CEA) [33]. Owing to its modular component structure, the APSS allows the user to efficiently implement and assess engine cycles with different levels of detail and technology.

In the case of the present paper, APSS was used to create a feedforward neural network engine surrogate model [19]. This was done to accommodate the varying engine design thrust and subsequent sizing during the BLADE aircraft design loop while simultaneously decreasing the overall calculation runtime. The implemented engine surrogate model allows the user to vary key design parameters (e.g., design streamtube net thrust or design altitude) alongside a variety of off-design inputs (e.g., flight Mach number, altitude, thrust lever angle), resulting in a wide array of engine designs and covering the full mission envelope of the associated aircraft design.

2.3. Previous Relevant Studies in Literature

Similar studies have been conducted on the integration and assessment of a BLI fan at the rear fuselage and its assessment on the overall aircraft level. One such project is Distributed Propulsion and Ultra-high By-Pass Rotor Study at Aircraft Level (DisPURSAL), which investigated the effect of BLI using distributed propulsion, as well as a single-fan concept for an Airbus A330-class aircraft [34]. Another project on the aircraft-level investigation of the propulsive fuselage concept is the Concept Validation Study for Fuselage Wake-Filling Propulsion Integration (CENTRELINE) [35]. This project also had an Airbus A330-class aircraft as its research focus, further investigating the single-BLI concept. However, these projects considered wide-body aircraft as the aircraft class to be investigated. Other relevant aircraft concepts with BLI propulsion are the Boeing SUGAR Freeze [36] and NASA STARC-ABL [37,38], both of which investigated narrowbody aircraft with BLI propulsion.

To determine the effect of the propulsion technology on the overall aircraft level, a conventional aircraft with a technology projection into the entry-into-service year of each concept aircraft is required. Therefore, a reference aircraft representing the state-of-the-art, alongside a baseline aircraft representing the technology level in the year in which the concept aircraft with BLI propulsion will enter service, was defined in each project. This practice is widely employed in other research projects, including the assessment of a technology brick at the overall aircraft level [39]. Examples of other baseline aircraft are the AVACON research baseline [40] and DLR-F25 [41].

2.4. Top-Level Aircraft Requirements

The first step of this study is the definition of TLARs, which the aircraft must fulfill since the airframe, engine, and aircraft subsystems are sized accordingly. For the HOPE-BL, a set of TLARs appropriate for a short-to-medium-range (SMR) aircraft in the A320neo-class with an EIS in 2035 was determined based on the state-of-the-art values of the A320neo [6] and an expert-level workshop conducted within the project. These values are used throughout the project timeline and are listed in

Table 1. Top-level aircraft requirements.

Requirement	Value
Entry into service year	2035
Design range	3000 nmi
Design the number of passengers (single-class)	180
Design payload	18,000 kg
Maximum payload	19,260 kg
Maximum operating Mach number	0.82
Initial cruise altitude	≥FL330
One-engine-inoperative ceiling altitude	FL170
Maximum operating altitude	FL410 *
Take-off field length (MTOM, ISA, SL)	≤2000 m
Landing field length (MLM, ISA, SL)	≤1900 m
Second segment climb	18,000 kg, DEN, ISA + 20 K ≥1000 nmi range
Time-to-climb	≤25 minutes *
Airport compatibility	ICAO Code C Wingspan < 36 m

* Slight variations (10–15%) are acceptable for these requirements.

.

Most of the TLARs are in agreement with the HOPE-R, such as the number of passengers in the single-class configuration, initial cruise altitude, and maximum payload capacity. Nevertheless, some other parameters have been changed to reflect conditions in 2035, such as an increase in the mass per passenger from 95 kg to 100 kg, defining a design payload of 18,000 kg instead of 17,100 kg. The maximum operating altitude has been increased to FL410 in accordance with the expected optimal trajectories that will be flown by the aircraft. The design mission block range has also been increased from 2950 nmi to 3000 nmi for the HOPE-BL. Some additional criteria have been added to the TLARs list, such as the mission to evaluate the second-segment climb gradient after take-off and the one-engine-inoperative (OEI) ceiling altitude, which are handled as off-design checks for the HOPE-R but may be the engine sizing case for the HOPE-BL.

In addition to the TLARs, several parameters regarding the design and typical missions must be defined to establish a consistent set of sizing and evaluation rules for all aircraft. For both missions, a climb schedule of 250 knots calibrated airspeed (KCAS) under FL100, 300 KCAS above FL100, and Mach 0.78 after reaching the crossover altitude is used. Both missions use a cruise Mach number of 0.78. The reserve mission segment, which consists of a 200 nmi diversion at FL250 and Mach 0.65, 30 minutes of holding at 1500 ft and Mach 0.38 as the final reserve, and 3% of the trip fuel as the contingency, is also kept constant for both missions. The International Standard Atmosphere (ISA) deviation for the design mission is set to ISA + 10 K to provide a buffer of excess performance when the aircraft operates under off-design conditions. Therefore, no ISA deviation is applied to the typical mission, since no such buffer is required. Additionally, the typical mission features a stage length of 800 nmi, which is flown much more often and therefore is more representative for comparing mission block fuel burn regarding the relevant operational range [42,43,44], in contrast to the design range of 3000 nmi.

2.5. Technology Assumptions

The following technology assumptions represent the improvements in certain components of the aircraft compared with the state-of-the-art. Such improvements can be expressed in terms of the mass, performance, or design parameters of each component. For the first two cases, the improvement compared to the state-of-the-art is expressed as a factor between 0 and 1, where 1 represents no improvement. These factors are applied similarly to an additional calibration factor for mass or performance estimation. A list of the applied technology assumptions for the HOPE-BL is presented in

Table 2. Technology assumptions.

Parameter	Value	Source
Airframe structural component masses, except engine pylons	0.9	[45], BHL estimation
Engine pylon mass	1.0	BHL estimation
Wing aspect ratio	≤12	BHL estimation
Wave drag airfoil technology factor	0.95	[25]
Fly-by-light flight control system mass	0.756	[46,47]
Hydraulic system mass	0	[46]
Environmental control system mass	1.7	[46]
Electric system mass	1.73	[46]
Operator items & furnishings mass	0.9	[46,48], BHL estimation

. These factors are based on the cited sources and an expert workshop, where each value was discussed and adjusted accordingly.

Most of the technology factors are related to mass estimation. The structural components of the airframe, except for the engine pylons, are assumed to be 10% lighter than the state-of-the-art, owing to improvements in manufacturing techniques and the use of advanced materials such as carbon-fiber-reinforced polymer (CFRP). The same reduction is also forecast for the operator items and furnishings masses for the same reasons as above. No improvement is assumed for the engine pylon masses.

The HOPE-BL employs an all-electric subsystem architecture. This has a considerable influence on the technology factors for the subsystem masses. The hydraulic system is replaced by a significantly heavier electric system, which increases the mass by 73%. The mass of the environmental control system also increases by 70%. Although the flight control system is also affected by the elimination of the hydraulic system, the use of a fly-by-light architecture [47] allows an all-electric flight control system that is 24.4% lighter than the state-of-the-art.

The maximum wing aspect ratio, which is one of the major parameters influencing the induced drag calculation (as showcased in Appendix F of [11]) and is the main driver for the corresponding aerodynamic improvements presented in Section 4.2, is set to 12. The wave drag airfoil technology factor, which influences the drag divergence Mach number, is described in [25] as a parameter that indicates how well an airfoil is designed for supercritical conditions. It is set to the highest possible value, yielding the lowest wave drag for a given flight condition and geometric parameters, serving as the primary driver that leads to the wave drag reductions shown in Section 4.2.

2.6. Aircraft Sizing Heuristics

The determination of the wing area and the engine design thrust in conceptual aircraft design is fundamental to the resulting aircraft. These two parameters significantly influence the size of all other components and systems, and determine the amount of available lift and excess thrust, which are the primary parameters affecting aircraft performance. Therefore, the relevant TLARs have been categorized into two groups. After checking all the criteria, the wing area and design thrust are adapted according to the most stringent criterion. The requirements used are listed in

Table 3. Wing and engine sizing requirements.

Requirement	Value	Group
Approach speed at maximum landing mass	<140 KCAS	Wing
Payload-range capability	3420 nmi, 15,120 kg	Wing
Take-off field length (MTOM, ISA, SL)	≤2000 m	Engine
Second segment climb gradient (18,000 kg, DEN, ISA + 20 K, ≥1000 nmi range)	>2.4% <3657 m TOFL	Engine
OEI ceiling altitude SEP (ISA + 10, 95% MTOM)	>300 fpm	Engine
Service ceiling altitude SEP (ISA, 82% MTOM)	>300 fpm	Engine
Top-of-climb SEP (ISA + 10 K, TOC mass)	>300 fpm	Engine

.

The HOPE-BL is designed to replace the A320neo-class HOPE-R; thus, it should be designed to achieve a similar payload-range performance. This ensures that both aircraft can cover the same routes. The critical point in the payload-range diagram of the HOPE-R for the sizing of the HOPE-BL was determined to be the point at maximum fuel capacity and the corresponding payload to achieve MTOM, which is at 3420 nmi range and 15,120 kg payload. The wing size affects the achievable range at this point, since it determines the maximum fuel capacity, resulting in an indirect minimum wing size requirement.

3. Engine Design and Results

The baseline aircraft for the HOPE project is powered by two under-wing-mounted Ultra-High Bypass Ratio (UHBR) Geared Turbofan (GTF) engines. The engine architecture features a bypass and a core stream. The compression section of the baseline engine consists of a fan, a three-stage Intermediate-Pressure Compressor (IPC), and an eight-stage High-Pressure Compressor (HPC). The HPC is driven by a corresponding two-stage High-Pressure Turbine (HPT), while the fan and IPC power are extracted from a four-stage Low-Pressure Turbine (LPT). As required for a GTF engine, the fan is linked to the low-pressure shaft via a gearbox with a gear ratio of 3.72. The cooling air for both the HPT and LPT is extracted from the HPC. To operate the cycle in part power without exceeding the compressor surge margins, a handling bleed from the IPC can be routed to the bypass duct. An overview of the engine architecture is shown in Figure 2.

The baseline engine model is fueled by Jet-A1 with a specified lower heating value of 42.8 MJ/kg. The design point of the engine is set to TOC conditions with an overall pressure ratio of 50 and a burner exit temperature (T₄) of 1725 K. As the aircraft is equipped with an all-electric subsystem architecture, 175 kW of high-pressure shaft power is extracted during all flight phases. In combination with the required design streamtube thrust of 19.8 kN, a TSFC of 14.25 g/s/kN for TOC conditions and 13.85 g/s/kN for cruise conditions is achieved. Compared to the GTF model used in HOPE-R with a TSFC of 14.8 g/s/kN at a representative cruise point, an improvement of 6.4% is reached.

Table 4. Baseline engine parameters for the Top-of-Climb, cruise, and take-off conditions.

Parameter	Unit	Top-of-Climb	Cruise	Take-Off
Mach number	[-]	0.78	0.78	0.23
Altitude	[m]	10,668	10,668	0
ISA temperature deviation	[K]	10	0	0
Streamtube thrust	[kN]	19.8	17.0	94.2
Net specific thrust	[m/s]	86.8	74.9	164.9
Streamtube TSFC	[g/s/kN]	14.25	13.85	9.02
Bypass ratio	[-]	15.6	16.2	14.5
Overall pressure ratio	[-]	50.0	45.6	48.0
Inlet mass flow (w2)	[kg/s]	231	229	578
Compressor exit temperature (T3)	[K]	865	796	953
Burner exit temperature (T4)	[K]	1725	1568	1885
Fuel-air ratio	[-]	0.027	0.023	0.030
Power offtake	[kW]	175	175	175

presents an overview of the key operating points for TOC conditions, cruise, and take-off at the rotation point, for the baseline engine.

Additionally, in order to avoid exceeding the operational limits of the engine, separate ratings for take-off and maximum climb (i.e., remaining mission excluding take-off) were implemented, taking into account the maximum allowable material temperatures and critical spool speeds. A visualization of the maximum climb-rated streamtube thrust with the corresponding TSFC values is shown in Figure 3.

4. Aircraft Design and Results

This chapter presents the synthesis of all technology assumptions as well as the engine design for the overall aircraft. The HOPE-BL and HOPE-R are compared with regard to all relevant disciplines: Design sensitivities for drag, fuel flow, and empty mass are shown to illustrate the aircraft’s response to small changes.

4.1. Geometries

A three-view image of the HOPE-BL is shown in Figure 4. Additional geometric parameters of the aircraft, compared with the values of the HOPE-R, are listed in

Table 5. Geometric parameters.

Parameter	Unit	HOPE-R	HOPE-BL
Wing area	m2	128.0	111.1
Wing aspect ratio	-	9.9	11.7
Wing mean aerodynamic chord	m	4.28	3.75
Main landing gear length	m	2.87	3.03

. One significant change in the geometric parameters is the wing area, which shrinks by 13.2% compared to the HOPE-R, which is caused by the overall mass reduction. The 36-m-airport-gate constraint is fulfilled by limiting the wingspan, as the wing can attain an aspect ratio of 11.7, which is very close to the selected maximum of 12 while also being span-limited. Therefore, the winglets are retained, instead of employing folding wingtip devices to increase the wingspan.

4.2. Aerodynamics

The aerodynamic performance of the HOPE-BL and HOPE-R under equal conditions, i.e., for the same Mach number, C_L, and altitude, which are representative of the HOPE-R, are listed in

Table 6. Aerodynamic performance.

Parameter	Unit	HOPE-R	HOPE-BL
Altitude	ft	35,000	35,000
Mach number	-	0.78	0.78
Lift coefficient (CL)	-	0.57	0.57
Drag coefficient (CD)	-	0.0316	0.0291
Lift-to-drag ratio (L/D)	-	18.03	19.57
Zero-lift drag coefficient (CD0)	-	0.0172	0.0180
Induced drag coefficient (CDi)	-	0.0112	0.0091
Wave drag coefficient (CDw)	-	0.0020	0.0011
Miscellaneous drag coefficient	-	0.0012	0.0009

.

Under equal conditions, the HOPE-BL is aerodynamically 8.5% more efficient than the HOPE-R. A significant part of this efficiency gain can be attributed to the increase in the aspect ratio to 11.7, which leads to a reduction in the induced drag coefficient of approximately 19%. Additionally, the selected wave-drag airfoil technology factor leads to a 45% decrease in wave drag. The increase in the zero-lift drag coefficient for the HOPE-BL is explained by the fact that this value is normalized to the wing area of each aircraft, and the total parasite drag area decreases overall for the HOPE-BL, which matches expectations.

In addition to comparing the aerodynamic performance at a given constant lift coefficient, it is also important to analyze the aerodynamic performance at a representative operating condition of both aircraft and over a range of lift coefficients, i.e., comparing drag polars. Such a comparison plot is shown in Figure 5.

This figure indicates the overall improvement in aerodynamic efficiency between the HOPE-R and HOPE-BL, which was already presented in the preceding point-based comparison. An interesting observation is that the optimum lift coefficient for the HOPE-BL is 10% higher than that for the HOPE-R (0.64 and 0.58, respectively). This is explained by the reduced impact of induced drag, which is driven by the increase in aspect ratio, as well as the significant reduction in wave drag based on the corresponding technology assumption. Furthermore, both aircraft operate very close to their optimal aerodynamic efficiencies at their average lift coefficients during the typical mission, as can be seen from the graph.

4.3. Masses

The design masses of the HOPE-BL and the HOPE-R are listed in

Table 7. Mass properties of HOPE-R and HOPE-BL.

Parameter	Unit	HOPE-R	HOPE-BL	Δ
Maximum take-off mass	kg	79,000	69,990	−11.4%
Maximum landing mass	kg	67,400	60,191	−10.7%
Operating empty mass	kg	44,925	38,787	−13.6%
Maximum fuel mass	kg	18,729	14,131	−24.6%

. The design masses for the HOPE-R are based on [6], whereas the component mass breakdown is based on [8]. As expected, the MTOM of the HOPE-BL is reduced by 11.4% because of the 13.6% reduction in OEM and further reductions in the design fuel, which is explained further in Section 4.4. The maximum fuel mass shrinks considerably because of the decrease in the wing area. A more detailed component mass breakdown for both aircraft is provided in Appendix A.

4.4. Performance

Before evaluating the mission performance of a proposed aircraft, it is important to ensure that all the design requirements are met.

Table 8. Performance requirements and results for the HOPE-BL.

Parameter	Requirement	HOPE-BL
Approach speed at maximum landing mass	<140 KCAS	128 KCAS
Take-off field length (MTOM, ISA, SL)	≤2000 m	1935 m
Second segment climb gradient (18,000 kg, DEN, ISA + 20, ≥1000 nmi range)	>2.4% <3657 m TOFL	4.1% 2795 m TOFL 1500 nmi range
Time-to-climb *	25 min	28.2 min
OEI ceiling altitude SEP (ISA + 10 K, 95% MTOM)	>300 fpm	300 fpm
Service ceiling altitude SEP (ISA, 82% MTOM)	>300 fpm	500 fpm
TOC SEP (ISA + 10 K, TOC mass)	>300 fpm	346 fpm

* Not a strict requirement.

shows the performance requirements and results for the HOPE-BL.

Summarizing Table 8, the HOPE-BL meets or exceeds all performance requirements. It is also evident from this table that the OEI ceiling altitude SEP requirement was the relevant engine-sizing case. Even though the climb time requirement was violated, this is deemed acceptable since the deviation is not large enough (see Table 1).

Regarding the mission performance metrics, the block fuel and energy values for the HOPE-R and HOPE-BL for both the design and typical missions are listed in

Table 9. Mission performance of HOPE-R and HOPE-BL.

Mission	Unit	HOPE-R	HOPE-BL	Δ
Design mission * block fuel	kg	14,338	11,276	−21.4%
Typical mission ** block fuel	kg	4013	3176	−20.9%
Design mission * block energy per kg km	kJ/(kg km)	6.57	4.83	−26.5%
Typical mission ** block energy per kg km	kJ/(kg km)	6.78	5.1	−24.8%

* HOPE-R design mission: 17,100 kg, 2950 nmi. HOPE-BL design mission: 18,000 kg, 3000 nmi. ** HOPE-R typical mission: 17,100 kg, 800 nmi. Typical HOPE-BL mission: 18,000 kg, 800 nmi.

. Even though the HOPE-BL is sized for ISA + 10 K conditions during the design mission, the presented values are given in ISA conditions for all missions to ensure comparability.

The synthesis of the effects regarding the technology assumptions in aerodynamics and masses, along with the improved engine, leads to a decrease of 21.4% in the design mission and 20.9% in the typical mission block fuel consumption for the HOPE-BL. Because of the increase in payload capacity for both missions and the 50 nmi increase in the design range, the HOPE-BL shows an even larger decrease in the energy consumption per kg of payload carried and per km of distance flown.

The payload-range performance of HOPE-BL and HOPE-R are shown in Figure 6. The design goal of covering the entire operational envelope of the HOPE-R with the HOPE-BL has been achieved, as almost all points on the blue line are either below or matching the orange line, with only a slight deviation around 3500 nmi. This graph also provides insight into why the approach speed at the maximum landing mass for the HOPE-BL is not the criterion for wing area sizing. It would be possible to further reduce the wing size and still satisfy the approach speed requirement; however, the available fuel volume inside the wing would then decrease in a manner such that the entire operational envelope of the HOPE-R would not be covered. Additional measures, such as planform variations or fuel tanks inside the fuselage, can be employed to increase the maximum fuel volume. However, these measures were not further explored in this study.

4.5. Design Sensitivities

Determining design sensitivities as part of an aircraft design study is essential to understand the aircraft model and its response to small changes in important design parameters, as well as to derive linear factors to approximate the impact of future incremental changes to the design. An example of such a change is the integration of a new engine into the HOPE-BL while retaining the overall aircraft configuration. Therefore, a sensitivity study was conducted for the HOPE-BL, where mass, drag, and fuel flow were evaluated as different parameters. To calculate the sensitivities for these parameters, the corresponding factors or increments are applied to the aircraft, and the aircraft is then redesigned to incorporate the effects of these changes on the overall aircraft level, including snowball effects. Each category was analyzed using an aircraft pair covering both positive and negative increments and averaging the results from both. The change in drag was applied as an additional factor, similar to a calibration factor, in addition to the existing aerodynamic calibration. The fuel flow study was also set up in a similar manner by applying an additional fuel flow factor to the engine model. For the mass increment, an additional mass was added on top of the OEM at the same CG. The resulting data are shown in Figure 7.

The drag increment has a substantial effect on the TOC thrust since any increase in drag is directly translated into a higher thrust demand. This leads to a much larger engine size, which leads to an increase in OEM of almost twice as much as the fuel flow increment study, which is also reflected in the MTOM sensitivity. The block fuel sensitivities in both studies are comparatively similar. An interesting result from the OEM increment study is the 1.4% resulting OEM increase from a 1% increment, showing a snowball effect of 40%. Other aspects of this study are in accordance with the expected values and do not warrant further explanation.

5. Summary and Conclusions

To investigate and assess the impact of the HOPE technology bricks, such as the BLI propulsor or fuel cell auxiliary power unit, the definition a comparable baseline aircraft with the same technology level is required. In this context, the TLARs and technology assumptions for 2035 are specified. Using these assumptions and the aircraft design environment BLADE, the HOPE-BL is conceptually defined. The HOPE-R, with a technology level corresponding to an EIS in 2015, is used to assess the performance improvement of the HOPE-BL.

Compared to the HOPE-R, the HOPE-BL has a 13.6% reduced OEM, which results mainly from the structural technology assumptions, leading to an 11.4% reduced MTOM and 5.3% increased design payload capacity, originating from the increase in the mass per passenger. The aerodynamic efficiency is increased by 8.5%, owing to the increase in the aspect ratio and a substantial decrease in the wave drag due to a technological assumption. The GTF engine used in the HOPE-R is replaced with a UHBR GTF engine, which is 6.4% more efficient in terms of TSFC at a corresponding representative cruise point for both aircraft. When integrated, all these effects yield a reduction in block fuel of 21.4% for the design mission and 20.9% for the typical mission between HOPE-BL and HOPE-R. The selected design sensitivities of HOPE-BL are also shown to facilitate the assessment of small changes in the design in the future.

In conclusion, the HOPE-BL, as presented in this paper, is a consistent and plausible dataset for an SMR aircraft with an EIS in 2035, enabling the integration of further technology bricks. The presented design sensitivities enable the evaluation of further technologies without them being explicitly modeled, making this dataset more useful for all researchers.